A method for ground-ground correlation of hypersonic boundary layer transition data

Abstract

The present invention discloses a method for the sky-ground correlation of hypersonic boundary layer transition data, which firstly uses the difference between the noise level of the incoming flow under the hypersonic wind tunnel and the actual flight conditions to calculate the boundary layer thickness difference; then uses the boundary layer thickness difference to correct the wind tunnel incoming flow unit Reynolds number; finally uses the corrected incoming flow unit Reynolds number to calculate the predicted value of the boundary layer transition Reynolds number, thereby achieving the purpose of predicting the boundary layer transition under the actual flight conditions based on the wind tunnel test data. The present invention is a method for predicting the boundary layer transition under the actual flight conditions based on the ground wind tunnel test data in the field of hypersonic boundary layer transition, which is simple to operate and has high prediction accuracy.

Description

Technical Field

[0001] The invention relates to a method for ground-to-air correlation of hypersonic boundary layer transition data, and belongs to the field of hypersonic aerodynamics. Background Art

[0002] Wind tunnel testing is one of the main means of predicting hypersonic boundary layer transition. However, due to the limitation of ground wind tunnel equipment, it is often impossible to fully simulate the real flight environment, so that the transition data of ground wind tunnel tests are not completely consistent with the transition conditions under real flight conditions. Among them, the incoming noise level is one of the main factors causing the difference between the earth and the sky in boundary layer transition data. At present, there is still a lack of effective earth-sky correlation methods for hypersonic boundary layer transition data, which can accurately predict the boundary layer transition under real flight conditions based on the transition data of ground wind tunnel tests. Summary of the invention

[0003] The purpose of the present invention is to overcome the above-mentioned defects and provide a method for the sky-ground correlation of hypersonic boundary layer transition data. First, the difference in the noise level of the incoming flow under the hypersonic wind tunnel and the actual flight conditions is used to calculate the boundary layer thickness difference; then the boundary layer thickness difference is used to correct the wind tunnel incoming flow unit Reynolds number; finally, the corrected incoming flow unit Reynolds number is used to calculate the predicted value of the boundary layer transition Reynolds number, thereby achieving the purpose of predicting the boundary layer transition under actual flight conditions based on wind tunnel test data. The present invention is a method for predicting the boundary layer transition under actual flight conditions based on ground wind tunnel test data in the field of hypersonic boundary layer transition. The method is simple to operate and has high prediction accuracy.

[0004] In order to achieve the above-mentioned object of the invention, the present invention provides the following technical solutions:

[0005] A method for space-ground correlation of hypersonic boundary layer transition data comprises the following steps:

[0006] (1) The difference between the incoming flow parameters under wind tunnel test conditions and those under actual flight conditions is less than 20% of the incoming flow parameters under actual flight conditions; the leading edge radius of the ground wind tunnel test model is consistent with that of the actual aircraft; the attitude angle difference is less than 1°; and the surface roughness difference is less than 20% of the surface roughness of the actual aircraft;

[0007] (2) calculating a difference Δδ between the boundary layer thickness under the wind tunnel test condition and the boundary layer thickness under the actual flight condition according to the difference ΔNL between the onflow noise level under the wind tunnel test condition and the actual flight condition; the boundary layer is a laminar boundary layer;

[0008] (3) The boundary layer thickness difference Δδ is used to correct the incoming flow unit Reynolds number Re under wind tunnel test conditions. ∞g , get the corrected incoming flow unit Reynolds number Re' ∞g ;

[0009] (4) Establishing the boundary layer transition Reynolds number Re T and the incoming flow unit Reynolds number Re ∞ Calculation formula, substitute the modified incoming flow unit Reynolds number Re' ∞g , the predicted value of the boundary layer transition Reynolds number Re under the actual flight conditions predicted by the wind tunnel test is obtained Tf .

[0010] Furthermore, the incoming noise level is the Pitot pressure pulsation Average Pitot pressure ratio.

[0011] Furthermore, in step (1), the incoming flow parameters include the incoming flow Mach number Ma and the incoming flow unit Reynolds number Re∞.

[0012] Furthermore, under actual flight conditions, the speed of the aircraft is ≥ 5 Mach.

[0013] Furthermore, in the step (2), the calculation formula Δδ=f(ΔNL) for calculating the difference Δδ between the boundary layer thickness under wind tunnel test conditions and the boundary layer thickness under actual flight conditions based on the difference ΔNL in the incoming flow noise level is a linear formula, and the linear coefficient in the linear formula is obtained by fitting the test data under different noise level conditions of multiple hypersonic wind tunnels.

[0014] Furthermore, in step (3), the corrected unit Reynolds number of the incoming flow under wind tunnel test conditions is

[0015]

[0016] in, x is the flow distance from the measuring point to the front edge stationary point of the model, T w is the model wall temperature, T e is the total temperature of the incoming flow, P r is the Prandtl number, γ is the specific heat ratio of the incoming flow, M e is the Mach number at the outer edge of the boundary layer.

[0017] Furthermore, in step (4), the boundary layer transition Reynolds number ReT and the incoming flow unit Reynolds number are calculated as a linear formula Re T =f(Re ∞ ), the linear coefficients in the linear formula are obtained by fitting the test data under wind tunnel test conditions or under actual flight conditions, the test data under wind tunnel test conditions are measurement data of wind tunnel tests conducted under different noise level conditions using a series of test models with the same leading edge radius, that is, the leading edge radii of the test models in the wind tunnel tests are equal and equal to the leading edge radius of the actual aircraft.

[0018] Furthermore, in step (4), the linear formula Re T =f(Re ∞ ) is obtained by fitting the test data under real flight conditions, and the intercept is obtained by fitting the test data under wind tunnel test conditions;

[0019] The test data under real flight conditions include the unit Reynolds number and the transition Reynolds number of the incoming flow under real flight conditions; the test data under wind tunnel test conditions include the unit Reynolds number and the transition Reynolds number of the incoming flow under wind tunnel test conditions.

[0020] Compared with the prior art, the present invention has the following beneficial effects:

[0021] (1) In the method for the sky-ground correlation of hypersonic boundary layer transition data of the present invention, the influence of the incoming flow noise is taken into account, and the incoming flow unit Reynolds number under the wind tunnel test condition is corrected based on the difference between the boundary layer thickness under the wind tunnel test condition and the boundary layer thickness under the actual flight condition, so that it is closer to the data obtained under the actual flight condition;

[0022] (2) In a method for calculating the sky-ground correlation of hypersonic boundary layer transition data of the present invention, a calculation formula for the boundary layer transition Reynolds number and the incoming flow unit Reynolds number is established based on historical test data under wind tunnel test conditions and actual flight conditions, further improving the accuracy of the calculation results;

[0023] (3) The present invention establishes the relationship between the difference in the incoming flow noise level under wind tunnel test conditions and under actual flight conditions and the boundary layer transition Reynolds number under actual flight conditions through the difference in the boundary layer thickness under wind tunnel test conditions and the boundary layer thickness under actual flight conditions, and the incoming flow unit Reynolds number under wind tunnel test conditions, thereby realizing the boundary layer transition prediction under actual flight conditions of hypersonic aircraft;

[0024] (4) The present invention realizes the correction of the oncoming noise level, which is a key influencing factor of boundary layer transition, and achieves the purpose of predicting boundary layer transition under actual flight conditions based on transition data of ground wind tunnel tests. The method has the advantages of simple operation and high prediction accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 A schematic diagram of a method for ground-to-air correlation of hypersonic boundary layer transition data according to the present invention;

[0026] Figure 2 This is the transition prediction result of Example 1 of the present invention. DETAILED DESCRIPTION

[0027] The following detailed description of the present invention will make the features and advantages of the present invention more clear and explicit.

[0028] The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise noted.

[0029] like Figure 1 As shown, the present invention is a method for the earth-sky correlation of hypersonic boundary layer transition data. In the prediction of boundary layer transition under real flight conditions of hypersonic aircraft, according to the difference in the noise level of incoming flow under ground wind tunnel test conditions and real flight conditions, similar incoming flow states and similar conditions of model and aircraft states are selected, the difference in incoming flow noise levels is used to calculate the boundary layer thickness difference, the incoming flow unit Reynolds number is corrected using the boundary layer thickness difference, and the boundary layer transition Reynolds number is calculated using the corrected incoming flow unit Reynolds number, thereby achieving the purpose of predicting boundary layer transition under real flight conditions based on wind tunnel test transition data. The steps are as follows:

[0030] 1) Select a wind tunnel with airflow conditions close to those of real flight conditions and with similar conditions of the model and aircraft;

[0031] 2) The difference in the outflow noise level ΔNL between the hypersonic wind tunnel and the actual flight conditions is used to calculate the boundary layer thickness difference Δδ; the boundary layer thickness δ is the laminar boundary layer thickness.

[0032] 3) The boundary layer thickness difference Δδ is used to correct the incoming flow unit Reynolds number Re under wind tunnel test conditions ∞g ;

[0033] 4) Use the modified incoming flow unit Reynolds number Re' ∞g , calculate the boundary layer transition Reynolds number Re under real flight conditions Tf , thereby achieving the purpose of predicting the boundary layer transition under real flight conditions based on wind tunnel test data.

[0034] In step 1), the similarity of the incoming flow state refers to the incoming flow Mach number Ma and the incoming flow unit Reynolds number Re under the ground wind tunnel test conditions and the actual flight conditions. ∞ The difference is less than 20% of the corresponding parameters of the actual flight conditions; the similarity between the model and the aircraft state means that the leading edge radius r of the test model and the actual aircraft are consistent, the attitude angle difference is less than 1°, and the surface roughness Ra difference is less than 20% of the corresponding parameters of the aircraft.

[0035] In step 2), the incoming noise level (abbreviated as NL) is the Pitot pressure pulsation Average Pitot pressure The ratio of

[0036] The incoming noise level NL affects the boundary layer thickness δ. The second mode wave frequency of the model surface under different noise level conditions in multiple hypersonic wind tunnels is measured. Based on the relationship between the second mode wave frequency and the boundary layer thickness, the boundary layer thickness δ of the model surface under different incoming noise level conditions is converted. Using the same calculation grid and calculation method, the laminar flow field of the model surface under different wind tunnel incoming flow conditions is calculated to obtain the calculated value of the boundary layer thickness δ CFD The least squares method is used to fit the boundary layer thickness δ and the calculated boundary layer thickness δ CFD The difference δ-δ CFD The linear relationship between the boundary layer thickness δ and the wind tunnel inflow noise level NL is δ=f(NL).

[0037] In step 3), the boundary layer thickness δ is proportional to the square root of the incoming flow unit Reynolds number The inverse proportional relationship is obtained by correcting the incoming flow unit Reynolds number under the ground wind tunnel test conditions and the actual flight conditions by the difference in boundary layer thickness Δδ. ∞g , that is, the unit Reynolds number Re of the incoming flow under wind tunnel test conditions ∞g Predicted unit Reynolds number of incoming airflow under real flight conditions ∞f .

[0038] In step 4), the linear relationship between the transition Reynolds number and the incoming flow unit Reynolds number is obtained by fitting the historical test data. T =f(Re ∞ ), the historical test data can be either wind tunnel test data or flight test data;

[0039] The corrected incoming flow unit Reynolds number Re' ∞g Substitute into the relation Re T =f(Re ∞ ) is used to calculate the predicted value of the boundary layer transition Reynolds number under actual flight conditions. Tf .

[0040] Under real flight conditions, the aircraft speed is ≥ 5 Mach.

[0041] Example 1

[0042] The Mach number of the incoming flow in the FD-30 wind tunnel is Ma=6, and the unit Reynolds number of the incoming flow is Re ∞g =1.01E+07 / m, transition Reynolds number Re Tg=5.86E+06, the wind tunnel test transition data of the model leading edge radius r=2.5mm, using a hypersonic boundary layer transition data sky-ground correlation method of the present invention, the flight Mach number Ma of the HiFIRE-1 aircraft is predicted to be 5.31, the flight altitude H=21.03km, and the incoming flow unit Reynolds number Re ∞f =8.93E+06 / m, the ground wind tunnel test model is consistent with the leading edge radius of the aircraft, the unit Reynolds number and Mach number of the incoming flow under wind tunnel and real flight conditions are relatively close, and the wind tunnel test model is a cone model (half cone angle 7 degrees). The FD-30 wind tunnel transition test data and the HiFIRE-1 aircraft transition test data are both from the literature "Tu Guohua, Wan Bingbing, Chen Jianqiang, et al. Research on the stability of MF-1 blunt cone boundary layer and transition sky-sky correlation [J]. Chinese Science: Physics, Mechanics and Astronomy, 2019, 49: 124701".

[0043] It is assumed that the wind tunnel test flow conditions are consistent with the actual flight flow conditions, that is, it is assumed that the flow parameters under the wind tunnel test conditions are consistent with the flow parameters under the actual flight conditions, the ground wind tunnel test model is consistent with the state parameters of the aircraft, and the wind tunnel test flow noise level is NL g The noise level of the incoming flow under actual flight conditions is NL f ,but:

[0044] Boundary layer thickness under ground wind tunnel conditions:

[0045] δ g =δ CFDg -0.31179+16.181*NL g (1)

[0046] Boundary layer thickness under real flight conditions:

[0047] δ f =δ CFDf -0.31179+16.181*NL f (2)

[0048] δ CFDg and δ CFDf The calculated values ​​of laminar boundary layer thickness δ are respectively the calculated values ​​of laminar boundary layer thickness under ground wind tunnel conditions and under real flight conditions without considering the influence of incoming noise. CFD , δ CFDThe calculation method is as follows: the Reynolds average numerical simulation method is used, the same calculation grid and the same calculation method are used, the flow field around the model is calculated according to the actual incoming flow conditions, the velocity information is extracted along the normal direction of the wall surface at the model measuring point, and the velocity value when the velocity value no longer increases as the distance from the model wall gradually increases is selected as Ve, and the distance from the model wall surface at a velocity of 0.99Ve is selected as the calculated value of the boundary layer thickness δ CFD ; From the above calculation process, we can know that δ CFDg =δ CFDf .

[0049] Subtracting equation (2) from equation (1) yields the difference in boundary layer thickness between wind tunnel test conditions and actual flight conditions:

[0050] Δδ=δ g -δ f =16.181*(NL g -NL f )=0.629mm (3)

[0051] Among them, NL g and NL f They are the inflow noise levels under wind tunnel inflow conditions and real flight conditions. The inflow noise level under wind tunnel test conditions is calculated by measuring the pulsating pressure with a pitot pressure probe. The inflow noise level under real flight conditions cannot be measured directly. It is generally believed in the academic community that the inflow noise level under real flight conditions is about 0.01%;

[0052] The engineering calculation formula for boundary layer thickness is:

[0053]

[0054] make Where x is the flow distance from the measuring point to the front edge stationary point of the model, T w is the model wall temperature, T e is the total temperature of the incoming flow, P r is the Prandtl number, γ is the specific heat ratio of the incoming flow, M e is the Mach number at the outer edge of the boundary layer

[0055] The boundary layer thickness under ground wind tunnel conditions is:

[0056]

[0057] Boundary layer thickness under real flight conditions:

[0058]

[0059] Subtracting equation (6) from equation (5) yields the unit Reynolds number Re' corrected by the boundary layer thickness difference Δδ:∞g :

[0060]

[0061] The research results show that the slope of the linear fitting between the boundary layer transition Reynolds number and the incoming flow unit Reynolds number is related to the leading edge radius of the model. The slope of the HiFIRE-1 flight test transition data is fitted according to the test data points of the FD-30 wind tunnel (the incoming flow unit Reynolds number ∞g =1.01E+07 / m, transition Reynolds number Re Tg =5.86E+06) to calculate the intercept, and obtain the linear relationship between the boundary layer transition Reynolds number and the incoming flow unit Reynolds number:

[0062] Re T =2.8444*10 6 +0.29857*Re ∞ (8)

[0063] The corrected incoming flow unit Reynolds number Re' ∞g =2.64E+07 / m Substituting into the above formula (8), the predicted transition Reynolds number Re can be calculated. Tf =1.07E+07, which is consistent with the flight test transition Reynolds number of 1.02E+07 in the reference. Figure 2 Schematic diagram of the transition prediction results of this embodiment, which shows the establishment process of formula (8) and the prediction process of the boundary layer transition Reynolds number under flight conditions, and also shows that the results of this embodiment are in good agreement with the flight test transition Reynolds number in the reference.

[0064] The present invention has been described in detail above in conjunction with specific implementations and exemplary examples, but these descriptions cannot be understood as limiting the present invention. Those skilled in the art understand that, without departing from the spirit and scope of the present invention, a variety of equivalent substitutions, modifications or improvements may be made to the technical solution of the present invention and its implementation methods, all of which fall within the scope of the present invention. The scope of protection of the present invention shall be subject to the attached claims.

[0065] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for space-ground correlation of hypersonic boundary layer transition data, characterized in that: The following steps are involved: (1) The difference between the incoming flow parameters under wind tunnel test conditions and those under actual flight conditions is less than 20% of the incoming flow parameters under actual flight conditions; the leading edge radius of the ground wind tunnel test model is consistent with that of the actual aircraft; the attitude angle difference is less than 1°; and the surface roughness difference is less than 20% of the surface roughness of the actual aircraft; (2) calculating a difference Δδ between the boundary layer thickness under the wind tunnel test condition and the boundary layer thickness under the actual flight condition according to the difference ΔNL between the onflow noise level under the wind tunnel test condition and the actual flight condition; the boundary layer is a laminar boundary layer; (3) The boundary layer thickness difference Δδ is used to correct the incoming flow unit Reynolds number Re under wind tunnel test conditions. ∞g , get the corrected incoming flow unit Reynolds number Re' ∞g ; (4) Establishing the boundary layer transition Reynolds number Re T and the incoming flow unit Reynolds number Re ∞ Calculation formula, substitute the modified incoming flow unit Reynolds number Re' ∞g , the predicted value of the boundary layer transition Reynolds number under the actual flight conditions predicted by the wind tunnel test is obtained Tf ; In the step (3), the corrected unit Reynolds number of the incoming flow under the wind tunnel test conditions is in, x is the flow distance from the measuring point to the front edge stationary point of the model, T w is the model wall temperature, T e is the total temperature of the incoming flow, P r is the Prandtl number, γ is the specific heat ratio of the incoming flow, M e is the Mach number at the outer edge of the boundary layer.

2. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 1, characterized in that: The incoming noise level is the Pitot pressure pulsation Average Pitot pressure ratio.

3. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 1, characterized in that: In step (1), the incoming flow parameters include the incoming flow Mach number Ma and the incoming flow unit Reynolds number Re ∞ .

4. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 1, characterized in that: Under real flight conditions, the aircraft speed is ≥ 5 Mach.

5. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 1, characterized in that: In the step (2), the calculation formula Δδ=f(ΔNL) for calculating the difference Δδ between the boundary layer thickness under wind tunnel test conditions and the boundary layer thickness under actual flight conditions based on the difference ΔNL between the incoming noise levels is a linear formula, and the linear coefficient in the linear formula is obtained by fitting the wind tunnel test data under different noise level conditions.

6. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 1, characterized in that: In the step (4), the boundary layer transition Reynolds number Re T The calculation formula of the unit Reynolds number of the incoming flow is the linear formula Re T =f(Re ∞ ), the linear coefficients in the linear formula are obtained by fitting the test data under wind tunnel test conditions or under actual flight conditions, and the leading edge radius of the test model in the wind tunnel test is equal and equal to the leading edge radius of the actual aircraft.

7. The method for space-ground correlation of hypersonic boundary layer transition data according to claim 6, characterized in that: In step (4), the linear formula Re T =f(Re ∞ ) is obtained by fitting the test data under real flight conditions, and the intercept is calculated from the test data under wind tunnel test conditions; The test data under real flight conditions include the unit Reynolds number and the transition Reynolds number of the incoming flow under real flight conditions; the test data under wind tunnel test conditions include the unit Reynolds number and the transition Reynolds number of the incoming flow under wind tunnel test conditions.

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