Fairing pulsating pressure scaling wind tunnel test method for reducing space-ground difference

Abstract

The invention discloses a fairing pulsating pressure scaling wind tunnel test method capable of reducing space-ground difference. The method comprises the following steps: determining initial parameters; based on the determined initial parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test, arranging pulsating pressure measuring points, and carrying out transonic speed state fairing pulsating pressure environment measurement to obtain a measurement result; performing parameter optimization according to a comparison result of a measurement result and a standard parameter under a real flight test condition to obtain an optimal parameter of manual transition; and based on the optimal parameters, carrying out a transonic speed scaling manual transition pulsating pressure wind tunnel test to obtain test results of the configuration in different free incoming flow Mach number, attack angle and sideslip angle states. According to the method, the free incoming flow local shock wave-boundary layer interference state and the boundary layer separation state of the ground wind tunnel test and the flight test are ensured to be closer through a manual transition mode, and the ground test prediction precision of the carrier rocket fairing pulsating pressure environment is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of rocket environment design, and in particular relates to a fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth. Background Art

[0002] After the rocket takes off, it will experience subsonic, transonic, and supersonic flow states during its journey through the atmosphere. In areas where the cross-section changes sharply, such as the cone-column shoulder, the hammerhead contraction section, and the skirt-column front, unsteady flow phenomena such as shock wave-boundary layer interference, shock wave oscillation, detached flow separation, and attached turbulent pulsation will occur, resulting in unsteady pressure pulsations with significant nonlinear characteristics. The unsteady characteristics of the flow around the aircraft will cause large pressure pulsations in local areas of the aircraft surface, especially when the frequency range is close to the low-order mode, which may cause fatigue or even damage to the aircraft structure. At the same time, the pulsating pressure environment has a significant impact on the vibration and noise protection design of the internal instruments and equipment of the aircraft. Obtaining an accurate pulsating pressure environment of the launch vehicle is very important for the design of the full rocket load strength, dynamic characteristics, rocket body structure, and elastic stability.

[0003] Due to the unsteady characteristics of the pulsating pressure, it is currently difficult to accurately predict the pulsating pressure environment of the fairing through the CFD (Computational Fluid Dynamics) method. The ground transonic pulsating pressure wind tunnel test is still an important way to obtain the characteristics of the transonic pulsating pressure environment of the launch vehicle's ascent phase. In recent years, the launch vehicle engineering development units have found through regression analysis that there are very significant differences between the transonic pulsating pressure environment of multiple types of launch vehicles and the ground wind tunnel test. The root mean square pulsating pressure coefficient obtained by the pulsating pressure wind tunnel test is much higher than that of the flight telemetry, and there are also large differences in the pulsating pressure spectrum information.

[0004] With the continuous deepening of the refined design of the new generation of rockets, the demand for refined design of the fairing pulsating pressure environment is becoming increasingly urgent. Therefore, it is urgent to develop a fairing pulsating pressure scale wind tunnel test method that reduces the difference between heaven and earth, improve the ground test prediction accuracy of the launch vehicle fairing pulsating pressure environment, and support the smooth progress of the refined design of the whole rocket load strength, dynamic characteristics, rocket body structure, and elastic stability. Summary of the invention

[0005] The technology of the present invention solves the problem: overcomes the shortcomings of the prior art and provides a scaled-down wind tunnel test method for a fairing pulsating pressure that reduces the difference between heaven and earth, aiming to overcome the problem that there are significant differences between the existing wind tunnel test results of the pulsating pressure environment of launch vehicles and flight telemetry, and ensures that the local shock wave-boundary layer interference state and the boundary layer separation state of the free flow in the ground wind tunnel test and the flight test are closer through artificial transition, thereby improving the prediction accuracy of the ground test of the pulsating pressure environment of the launch vehicle fairing, and supporting the smooth progress of the refined design of the full rocket load strength, dynamic characteristics, rocket body structure, and elastic stability.

[0006] In order to solve the above technical problems, the present invention discloses a fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth, comprising:

[0007] Determine the initial parameters required for artificial transition;

[0008] Based on the determined initial parameters, a transonic scaled artificial transition pulsation pressure wind tunnel test was carried out for the launch vehicle aerodynamic configuration with fairing pulsation pressure flight telemetry data. Pulsation pressure measurement points were arranged along the fairing longitudinal section and typical circumferential section, and the fairing pulsation pressure environment in the transonic state was measured to obtain the measurement results.

[0009] Obtain standard parameters under real flight test conditions of launch vehicle aerodynamic configuration with fairing pulsation pressure flight telemetry data;

[0010] Compare the measurement results with standard parameters to obtain comparison results;

[0011] According to the comparison results, the initial parameters are adjusted and optimized to obtain the optimal parameters for artificial transition;

[0012] Based on the optimal parameters obtained, a transonic scaled-down artificial transition pulsation pressure wind tunnel test was carried out on the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data, and the test results of the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data under different free flow Mach numbers, angles of attack, and sideslip angles were obtained.

[0013] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth, the initial parameters include: the arrangement position of the roughness element, the height H of the roughness element, the diameter D of the roughness element and the diameter gap ratio λ of the roughness element.

[0014] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method to reduce the difference between heaven and earth,

[0015] The rough elements are arranged on the edge of the fairing ball head;

[0016] The height H of the rough element satisfies: k <H<δ; where H krepresents the critical height of roughness element transition, δ represents the thickness of local boundary layer;

[0017] The diameter D of the rough element is: D = D j 100; among them, D j Indicates the diameter of the arrow body;

[0018] The diameter-to-gap ratio λ of the rough element is: λ=0.3.

[0019] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference, the critical height H of the roughness element transition is determined by the following equations (1) to (4): k :

[0020]

[0021] Re x =ρμ x xμ···(2)

[0022] Re k =ρμ k H k μ···(3)

[0023]

[0024] Where x represents the axial distance of the rough element from the leading edge of the launch vehicle aerodynamic configuration with the flight telemetry data of the fairing pulsation pressure, Re x Reynolds number is the local Reynolds number. k represents the rough element Reynolds number, η k represents the transition dimensionless parameter, ρ represents the free flow density, μ x represents the local velocity at the outer edge of the boundary layer where the rough element is located, μ k represents the flow velocity at the top of the rough element, μ represents the free flow dynamic viscosity coefficient, c represents the viscosity proportional constant, T x represents the local temperature at the outer edge of the boundary layer where the rough element is located, T k Represents the top temperature of the rough element.

[0025] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference, the local boundary layer thickness δ is determined by the following formula (5):

[0026]

[0027] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth, the measurement results include: the root mean square pulsation pressure coefficient C p and the dimensionless power spectral density G(f).

[0028] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference, the root mean square pulsation pressure coefficient C at the measuring point is obtained by the following equations (6) to (8): p And the dimensionless power spectral density G(f):

[0029]

[0030]

[0031] Among them, P rms Indicates the pulsating pressure root mean square value, P LU represents the local unsteady pressure, P LS represents the time average of the local unsteady pressure, q represents the free flow pressure, Δf represents the frequency resolution, It represents the square of the instantaneous value of the pulsating pressure within Δf, t represents the sampling moment, and T represents the sampling time.

[0032] In the above-mentioned fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth, the standard parameters include: the standard value of the root mean square pulsation pressure coefficient of the measuring point C' p and the dimensionless power spectral density standard value G′(f).

[0033] In the fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference, the initial parameters are adjusted and optimized according to the comparison results to obtain the optimal parameters of the artificial transition, including: adjusting and optimizing the initial parameters according to the comparison results until C p With C p ′, as well as the difference between G(f) and G′(f) are all within the preset threshold range, and the parameters x′, H′, D′ and λ′ obtained at this time are used as the optimal parameters of artificial transition; wherein, x′ represents the optimal axial distance of the rough element from the leading edge point of the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data, H′ represents the optimal height of the rough element, D′ represents the optimal diameter of the rough element, and λ′ represents the optimal diameter-to-gap ratio of the rough element.

[0034] In the above-mentioned fairing pulsation pressure scaled wind tunnel test method that reduces the difference between heaven and earth, the test results of the aerodynamic configuration of the launch vehicle with the fairing pulsation pressure flight telemetry data under different free flow Mach numbers, angles of attack, and sideslip angles include: the root mean square pulsation pressure coefficient along the longitudinal section and typical circumferential section of the fairing under different free flow Mach numbers, angles of attack, and sideslip angles of the launch vehicle with the fairing pulsation pressure flight telemetry data, and the dimensionless power spectral density along the longitudinal section and typical circumferential section of the fairing.

[0035] The present invention has the following advantages:

[0036] The invention discloses a scaled-down wind tunnel test method for fairing pulsation pressure that reduces the difference between heaven and earth. The method ensures that the local shock wave-boundary layer interference state and boundary layer separation state of the free flow in the ground wind tunnel test and the flight test are closer through an artificial transition method, thereby improving the ground test prediction accuracy of the pulsation pressure environment of the launch vehicle fairing, and increasing the root mean square pulsation pressure coefficient of the cone shoulder of the fairing before and after the transition measure (such as Figure 5 As shown in the figure, it supports the smooth progress of the refined design of the entire arrow's load strength, dynamic characteristics, arrow body structure, and elastic stability. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Figure 1 It is a flow chart of steps of a fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference in an embodiment of the present invention;

[0038] Figure 2 It is an embodiment of the present invention. k and Schematic diagram of the relationship between ; Ma represents the flight Mach number;

[0039] Figure 3 1 is a schematic diagram of the difference between the root mean square pulsation pressure coefficient of a fairing and the earth in an embodiment of the present invention;

[0040] Figure 4 1 is a schematic diagram of the difference between the dimensionless power spectrum density of the fairing pulsation pressure and the sky and the earth in an embodiment of the present invention;

[0041] Figure 5 It is a schematic diagram comparing the root mean square pulsating pressure of the fairing shoulder with and without transition measures in an embodiment of the present invention. DETAILED DESCRIPTION

[0042] In order to make the objectives, technical solutions and advantages of the present invention more clear, the embodiments disclosed in the present invention will be further described in detail below with reference to the accompanying drawings.

[0043] One of the core ideas of the present invention is: in order to overcome the problem of significant differences between the earth and the sky between the results of the existing wind tunnel test of the pulsating pressure environment of the launch vehicle and the flight telemetry, a fairing pulsating pressure scaled wind tunnel test method is proposed to reduce the difference between the earth and the sky. Through artificial transition, the local shock wave-boundary layer interference state and the boundary layer separation state of the free flow in the ground wind tunnel test and the flight test are guaranteed to be closer, thereby improving the prediction accuracy of the ground test of the pulsating pressure environment of the launch vehicle fairing, and supporting the smooth progress of the refined design of the whole rocket load strength, dynamic characteristics, rocket body structure and elastic stability.

[0044] In this embodiment, the fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth mainly includes the following three aspects:

[0045] (I) Optimizing the artificial transition scheme for the launch vehicle pulsating pressure wind tunnel test

[0046] The Reynolds number of the ground wind tunnel test is lower than that of the actual flight state of the launch vehicle. When flying in the atmosphere, the boundary layer on the surface of the rocket body generally changes at the front cone of the fairing, which is close to the full turbulence state. In the ground wind tunnel test, the flow around the fairing will have a long part of the flow area that cannot fully simulate the flow state of the flight state. This will lead to significant changes in the turbulent flow state, shock wave-boundary layer interference and shock wave oscillation intensity at the shoulder of the fairing, which will lead to a significant difference between the pulsating pressure test results of the shoulder of the fairing cone and the flight telemetry results.

[0047] In order to reduce the impact of the difference in the free flow Reynolds number between the wind tunnel test and the real flight state on the pulsating pressure test results, an artificial transition method is adopted to obtain a fairing pulsating pressure environment with high Reynolds number characteristics. Commonly used artificial transition methods include but are not limited to local transition measures such as corundum, metal mixing wire and distributed columnar rough elements. Among them, the corundum rough belt has problems such as easy detachment of gravel under the influence of the flow, and the adhesive dosage and the width and height of the rough belt are difficult to accurately grasp through effective methods; the metal mixing wire has the problem of difficulty in ensuring the uniformity of the circumferential arrangement of the transition belt; after optimization, the distributed columnar rough element transition belt is adopted. This type of rough element transition belt controls the passive flow transition by controlling the height, shape, spacing, and arrangement of the distributed columnar rough elements. It has high transition efficiency, low additional friction resistance, and is easy to standardize and repeat the test. It is suitable for artificial transition in the transonic pulsating pressure wind tunnel test of launch vehicles.

[0048] (II) Calibration based on artificial transition pulsating pressure wind tunnel test method

[0049] Based on the optimal artificial transition method, a transonic scaled pulsation pressure wind tunnel test was carried out for the aerodynamic configuration of the launch vehicle with the flight telemetry data of the fairing pulsation pressure. The cylindrical rough element transition zone was arranged in the front cone area of ​​the fairing, and the pulsation pressure measurement points were arranged along the longitudinal section and the typical circumferential section of the fairing to measure the pulsation pressure environment of the fairing in the transonic state. After the test, the time domain pulsation pressure test data was processed to obtain the root mean square pulsation pressure coefficient and dimensionless power spectrum density of the measurement point. The flight test pulsation pressure measurement data was used as the baseline to correct the related parameters of the distributed cylindrical rough element transition zone (the arrangement position, height, diameter, and diameter gap ratio of the rough element). In order to reduce the difference between the earth and the sky in the pulsation pressure environment of the shoulder of the launch vehicle fairing, multiple rounds of comparative wind tunnel tests were carried out, and the transition zone rough element parameters that can significantly reduce the difference between the earth and the sky in the pulsation pressure environment of the fairing were obtained through optimization. The deviation between the root mean square pulsation pressure coefficient and power spectrum density of the ground wind tunnel test and the flight telemetry results is controlled within the acceptable threshold range, and the artificial transition pulsation pressure wind tunnel test method is calibrated.

[0050] (III) Based on the calibrated artificial transition pulsating pressure wind tunnel test method, the pulsating pressure environment of the fairing of the launch vehicle aerodynamic configuration with the flight telemetry data of the fairing pulsating pressure is obtained.

[0051] Based on the calibrated artificial transition pulsating pressure wind tunnel test method, a transonic pulsating pressure wind tunnel test is carried out for the aerodynamic configuration of the launch vehicle with fairing pulsating pressure flight telemetry data. The root mean square pulsating pressure coefficient and dimensionless power spectral density parameters along the longitudinal section and typical circumferential section of the fairing (typical circumferential sections include but are not limited to: fairing cone shoulder section and fairing inverted cone section) under different free flow Mach numbers, angles of attack, and sideslip angles are obtained to guide the design of the launch vehicle fairing pulsating pressure environment and support the refined design of the entire rocket load strength, dynamic characteristics, rocket body structure, and elastic stability.

[0052] Reference Figure 1 In this embodiment, the fairing pulsation pressure reduction wind tunnel test method for reducing the difference between heaven and earth includes:

[0053] Step 1: Determine the initial parameters required for artificial transition.

[0054] In this embodiment, the initial parameters mainly include:

[0055] Layout position of rough elements: In order to ensure sufficient transition of flow field turbulence near the shoulder of the rocket fairing, it is preferred to arrange the rough elements at the edge of the fairing ball head. If it is difficult to arrange the rough elements in this area (edge ​​of the fairing ball head), it is preferred to arrange the rough elements as close to the fairing ball head as possible.

[0056] The height H of the rough element: The height H of the rough element satisfies: H k <H<δ. Among them, H k represents the critical height of roughness element transition, and δ represents the thickness of the local boundary layer.

[0057] Rough element transition critical height H k It can be determined by the following formulas (1) to (4):

[0058]

[0059]

[0060] Re k =ρμ k H k μ···(3)

[0061]

[0062] Where x represents the axial distance of the rough element from the leading edge of the launch vehicle aerodynamic configuration with the flight telemetry data of the fairing pulsation pressure, Rex Reynolds number is the local Reynolds number. k represents the rough element Reynolds number, η k represents the transition dimensionless parameter, ρ represents the free flow density, μ x represents the local velocity at the outer edge of the boundary layer where the rough element is located, μ k represents the flow velocity at the top of the rough element, μ represents the free flow dynamic viscosity coefficient, c represents the viscosity proportional constant, T x represents the local temperature at the outer edge of the boundary layer where the rough element is located, T k represents the top temperature of the rough element. Formula (4) represents η k and The relationship between Figure 2 shown.

[0063] The local boundary layer thickness δ can be determined by the following formula (5):

[0064]

[0065] The diameter D of the rough element: D = D j 100; among them, D j Indicates the diameter of the arrow body.

[0066] The diameter-to-gap ratio λ of the roughness element is: λ=0.3.

[0067] Preferably, the rough element is arranged at the intersection of the fairing ball head and the front cone, and the height H of the rough element is equal to δ, D = D j 100, λ = 0.3, the fairing pulsation pressure reduction wind tunnel test method has the best effect on reducing the difference between the earth and the sky in the root mean square pulsation pressure environment of the launch vehicle fairing shoulder. Figure 5 shown.

[0068] Step 2: Based on the determined initial parameters, a transonic scaled artificial transition pulsating pressure wind tunnel test is carried out for the aerodynamic configuration of the launch vehicle with fairing pulsating pressure flight telemetry data. Pulsating pressure measuring points are arranged along the longitudinal section and typical circumferential section of the fairing, and the fairing pulsating pressure environment in the transonic state is measured to obtain the measurement results.

[0069] In this embodiment, the measurement results obtained in step 2 mainly include: the root mean square pulsation pressure coefficient C of the measuring point p And the dimensionless power spectrum density G(f) can be obtained by the following equations (6) to (8):

[0070]

[0071] Among them, P rms Indicates the root mean square value of pulsating pressure, P LU represents the local unsteady pressure, P LSrepresents the time average of the local unsteady pressure, q represents the free flow pressure, Δf represents the frequency resolution, It represents the square of the instantaneous value of the pulsating pressure within Δf, t represents the sampling moment, and T represents the sampling time.

[0072] Step 3, obtain the standard parameters of the launch vehicle aerodynamic configuration under the actual flight test conditions with the fairing pulsation pressure flight telemetry data.

[0073] In this embodiment, the standard parameters of the launch vehicle aerodynamic configuration under the actual flight test conditions with fairing pulsation pressure flight telemetry data mainly include: the standard value of the root mean square pulsation pressure coefficient C p ′ and the dimensionless power spectral density standard value G′(f).

[0074] Step 4: Compare the measurement result with the standard parameter to obtain a comparison result.

[0075] Step 5: According to the comparison results, the initial parameters are adjusted and optimized to obtain the optimal parameters for artificial transition.

[0076] In this embodiment, the initial parameters can be adjusted and optimized according to the comparison results between the measurement results and the standard parameters until C p With C p ′, and the difference between G(f) and G′(f) are all within the preset threshold range (such as Figure 3 and Figure 4 As shown in the figure, the parameters x', H', D' and λ' obtained at this time are used as the optimal parameters of artificial transition. Among them, x' represents the optimal axial distance between the rough element and the leading edge point of the aerodynamic configuration of the launch vehicle with the flight telemetry data of the fairing pulsating pressure, H' represents the optimal height of the rough element, D' represents the optimal diameter of the rough element, and λ' represents the optimal diameter-gap ratio of the rough element.

[0077] Preferably, the gradient descent method can be used for parameter optimization: x, H, D, λ in step 1 are used as initial parameters, and C p and C′ p The root mean square error of G(f) and the root mean square error of G′(f) are used as the objective function. The gradient of the objective function under the current parameter values ​​x, H, D, and λ is calculated, and the parameters are updated in the direction of the negative gradient until C p With C p ′, and the difference between G(f) and G′(f) are all within the preset threshold range, and the parameters x′, H′, D′ and λ′ obtained at this time are taken as the optimal parameters for artificial transition.

[0078] Step 6. Based on the obtained optimal parameters, a transonic-scaled artificial transition pulsation pressure wind tunnel test is carried out for the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data, and the test results of the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data under different free flow Mach numbers, angles of attack, and sideslip angles are obtained.

[0079] In this embodiment, the test results of the aerodynamic configuration of the launch vehicle with the flight telemetry data of the fairing pulsation pressure under different free flow Mach numbers, angles of attack, and sideslip angles mainly include: the root mean square pulsation pressure coefficient along the longitudinal section and typical circumferential section of the fairing under different free flow Mach numbers, angles of attack, and sideslip angles, and the dimensionless power spectral density along the longitudinal section and typical circumferential section of the fairing.

[0080] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

[0081] The contents not described in detail in the specification of the present invention belong to the common knowledge of the professionals in this field.

Claims

1. A wind tunnel test method for reducing the difference between heaven and earth for the pulsating pressure of a fairing, characterized in that: include: Determine the initial parameters required for artificial transition; Based on the determined initial parameters, a transonic scaled artificial transition pulsation pressure wind tunnel test was carried out for the launch vehicle aerodynamic configuration with fairing pulsation pressure flight telemetry data. Pulsation pressure measurement points were arranged along the fairing longitudinal section and typical circumferential section, and the fairing pulsation pressure environment in the transonic state was measured to obtain the measurement results. Obtain standard parameters under real flight test conditions of launch vehicle aerodynamic configuration with fairing pulsation pressure flight telemetry data; Compare the measurement results with standard parameters to obtain comparison results; According to the comparison results, the initial parameters are adjusted and optimized to obtain the optimal parameters for artificial transition; Based on the optimal parameters obtained, a transonic scaled-down artificial transition pulsation pressure wind tunnel test was carried out on the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data, and the test results of the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data under different free flow Mach numbers, angles of attack, and sideslip angles were obtained.

2. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 1 is characterized in that: The initial parameters include: the arrangement position of the rough element, the height H of the rough element, the diameter D of the rough element and the diameter-gap ratio λ of the rough element.

3. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 2 is characterized in that: The rough elements are arranged on the edge of the fairing ball head; The height H of the rough element satisfies: k <H<δ; where H k represents the critical height of roughness element transition, δ represents the thickness of local boundary layer; The diameter D of the rough element is: D = D j / 100; among them, D j Indicates the diameter of the arrow body; The diameter-to-gap ratio λ of the rough element is: λ=0.

3.

4. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 3 is characterized in that: The critical height H of the roughness element transition is determined by the following equations (1) to (4): k : Re x =rm x x / μ···(2) Re k =rm k H k / m···(3) Where x represents the axial distance of the rough element from the leading edge of the launch vehicle aerodynamic configuration with the flight telemetry data of the fairing pulsation pressure, Re x Reynolds number is the local Reynolds number. k represents the rough element Reynolds number, η k represents the transition dimensionless parameter, ρ represents the free flow density, μ x represents the local velocity at the outer edge of the boundary layer where the rough element is located, μ k represents the flow velocity at the top of the rough element, μ represents the free flow dynamic viscosity coefficient, c represents the viscosity proportional constant, T x represents the local temperature at the outer edge of the boundary layer where the rough element is located, T k Represents the top temperature of the rough element.

5. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 4 is characterized in that: The local boundary layer thickness δ is determined by the following equation (5):

6. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 4 is characterized in that: Measurement results, including: RMS pulsation pressure coefficient C at the measuring point p and the dimensionless power spectral density G(f).

7. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 6 is characterized in that: The root mean square pulsation pressure coefficient C of the measuring point is obtained by the following formulas (6) to (8): p And the dimensionless power spectral density G(f): Among them, P rms Indicates the root mean square value of pulsating pressure, P LU represents the local unsteady pressure, P LS represents the time average of the local unsteady pressure, q represents the free flow pressure, Δf represents the frequency resolution, It represents the square of the instantaneous value of the pulsating pressure within Δf, t represents the sampling moment, and T represents the sampling time.

8. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 7 is characterized in that: Standard parameters, including: Standard value of root mean square pulsation pressure coefficient C' at the measuring point p and the dimensionless power spectral density standard value G′(f).

9. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 8 is characterized in that: According to the comparison results, the initial parameters are adjusted and optimized to obtain the optimal parameters for artificial transition, including: according to the comparison results, the initial parameters are adjusted and optimized until C p With C p ′, as well as the difference between G(f) and G′(f) are all within the preset threshold range, and the parameters x′, H′, D′ and λ′ obtained at this time are used as the optimal parameters of artificial transition; wherein, x′ represents the optimal axial distance of the rough element from the leading edge point of the aerodynamic configuration of the launch vehicle with fairing pulsation pressure flight telemetry data, H′ represents the optimal height of the rough element, D′ represents the optimal diameter of the rough element, and λ′ represents the optimal diameter-to-gap ratio of the rough element.

10. The fairing pulsation pressure reduction wind tunnel test method for reducing the sky-ground difference according to claim 1, characterized in that: The test results of the aerodynamic configuration of the launch vehicle with flight telemetry data of fairing pulsation pressure under different free flow Mach numbers, angles of attack, and sideslip angles include: the root mean square pulsation pressure coefficient along the longitudinal section and typical circumferential section of the fairing under different free flow Mach numbers, angles of attack, and sideslip angles, and the dimensionless power spectral density along the longitudinal section and typical circumferential section of the fairing.

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