A method for calculating and correcting the linear load distribution of a launch vehicle

Abstract

A method for calculating and correcting the linear load distribution of a launch vehicle relates to the calculation and correction of the aerodynamic linear load distribution (the distribution density of the normal force coefficient along the axial direction) of the launch vehicle. Based on the numerical simulation results of the aerodynamic flow field of the launch vehicle, the aerodynamic linear load distribution is calculated by performing a segmented integration of the unit pressure and shear stress on the surface of the rocket body along the axial direction of the rocket body, and the aerodynamic linear load distribution is corrected by setting a target normal force coefficient and a target pressure center position, thereby solving the problem of large errors in the engineering calculation method. Compared with the traditional method that needs to distribute the aerodynamic force of each section of the launch vehicle according to the personal experience of the designer or test data, the present invention is based on the CFD numerical simulation results, reduces the dependence on the personal experience of engineers, and has high calculation efficiency and good robustness.

Description

Technical Field

[0001] The invention relates to a method for calculating and correcting the linear load distribution of a launch vehicle, and belongs to the field of launch vehicle load design. Background Art

[0002] For slender aircraft such as launch vehicles, it is necessary to calculate their aerodynamic line load distribution (the distribution density of the normal force coefficient along the axial direction) and correct the aerodynamic line load distribution according to the normal force coefficient of the entire rocket and the pressure center coefficient. The existing aerodynamic line load distribution calculation method is based on the engineering calculation method, which requires the aerodynamic force of each section of the launch vehicle to be distributed according to the designer's personal experience or test data. In the absence of test data, this method has a large error. Summary of the invention

[0003] The technical problem to be solved by the present invention is to overcome the shortcomings of the prior art and solve the problem of large calculation errors of aerodynamic line load distribution caused by engineering calculations and human factors.

[0004] The purpose of the present invention is achieved through the following technical solutions:

[0005] A method for calculating and correcting linear load distribution of a launch vehicle, comprising:

[0006] (1) Use CFD method to perform numerical simulation on the aerodynamic characteristics of the launch vehicle three-dimensional model, and output the aerodynamic characteristic parameters of the rocket body surface based on the numerical simulation results;

[0007] (2) Calculate the aerodynamic line load distribution using the piecewise integration method;

[0008] (3) Correcting the aerodynamic line load distribution according to the target normal force coefficient and the target pressure center position;

[0009] (4) Output the calculation results of aerodynamic line load distribution.

[0010] Preferably, the aerodynamic characteristic parameters of the rocket body surface at least include the central axial coordinate X of each grid unit on the rocket body surface. j 、Pressure P j , normal component of shear stress τ jy , unit area A j And the unit normal projection area A jy .

[0011] Preferably, the calculation method of the aerodynamic line load distribution includes:

[0012] (a) Set the aerodynamic line distribution load deployment node sequence x along the axial direction of the rocket body i , i=1,2,,N;

[0013] (b) According to the deployment node sequence xi Set the rocket body axial segment control volume Δx i , and calculate each segmented control volume Δx i The center coordinates s i :

[0014]

[0015]

[0016] (c) Axial segment control volume Δx of the rocket body i All CFD surface mesh elements within are integrated, and the segment normal force coefficient ΔC is calculated according to the following formula ni , where Q ref is the reference dynamic pressure, A ref The reference area of ​​the rocket is:

[0017]

[0018] (d) In the axial segment control volume Δx i The normal force coefficient along the axial distribution density C' is calculated according to the following formula ni , where C n is the normal force coefficient:

[0019]

[0020] Preferably, the distributed deployment node sequence x i The density is increased in areas with drastic flow changes and decreased in areas with gentle flow changes.

[0021] Preferably, for the satellite fairing nose cone, cone-column transition area, inverted cone, and tail wing installation area, the deployment nodes are encrypted, and the distance between two deployment nodes after encryption is not less than 3 times the local grid unit characteristic size; for the first-stage barrel section area, the nodes are sparsely deployed, and the distance between two deployment nodes after sparseness should not be greater than 20 times the local grid unit characteristic size.

[0022] Preferably, the aerodynamic line load distribution correction method includes:

[0023] (a) The normal force coefficient C of the rocket body can be calculated by the aerodynamic line load distribution n and the pressure center position X p :

[0024] C n =∑C′ ni Δx i

[0025] C n X p =∑C′ ni Δx i·s i

[0026] (b) Obtain the target normal force coefficient C of the rocket body through numerical simulation, wind tunnel test or deviation design n,t and target pressure center position X p,t , then there is a target aerodynamic line load distribution C′ ni,t satisfy:

[0027] C n,t =∑C′ ni,t Δx i

[0028] C n,t X p,t =∑C′ ni,t Δx i ·s i

[0029] (c) The corrected target aerodynamic line load distribution C′ can be calculated by the following formula: ni,t :

[0030] C′ ni,t =[k(xX p )+b]C′ ni

[0031] Where:

[0032]

[0033]

[0034] Preferably, the calculation results of the aerodynamic linear load distribution at least include the center coordinates s of each segment control body in the axial direction of the rocket body. i And the normal force coefficient along the axial distribution density value C' ni,t .

[0035] Preferably, the launch vehicle linear load distribution includes at least the axial distribution density of the normal force coefficient.

[0036] A computer-readable storage medium stores computer program instructions, which, when loaded and executed by a processor, enable the processor to execute a method for calculating and correcting the linear load distribution of a launch vehicle.

[0037] An electronic device, comprising:

[0038] Processor; and

[0039] a memory for storing computer program instructions;

[0040] When the computer program instructions are loaded and run by the processor, the processor executes a method for calculating and correcting the linear load distribution of a launch vehicle.

[0041] Compared with the prior art, the present invention has the following beneficial effects:

[0042] (1) Compared with the traditional engineering calculation method based on component combination, the linear load distribution of the launch vehicle calculated by the present invention based on the CFD numerical simulation results has higher accuracy;

[0043] (2) Compared with the traditional method that requires the aerodynamic force of each section of the launch vehicle to be distributed based on the designer's personal experience or test data, the present invention is based on CFD numerical simulation results, reduces the dependence on the engineer's personal experience, has high calculation efficiency and good robustness;

[0044] (3) Compared with the disadvantage of poor calculation accuracy of complex configuration launch vehicles by traditional methods, the present invention is not only applicable to single-core rockets, but also to complex configuration launch vehicles such as booster bundles, with high calculation accuracy and wide adaptability;

[0045] (4) Compared with the traditional method that cannot give the details of the local linear load distribution, the present invention can more accurately and reasonably give the local aerodynamic linear load distribution on each section by controlling the density of the node sequence. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] Figure 1 A schematic diagram of the aerodynamic simulation coordinate system provided by the present invention;

[0047] Figure 2 A schematic diagram of an axial segmented control body provided by the present invention;

[0048] Figure 3 A schematic diagram of the integral of the surface pressure and shear stress of the arrow body provided by the present invention;

[0049] Figure 4 The present invention provides aerodynamic linear load distribution of a launch vehicle. DETAILED DESCRIPTION

[0050] To make the objectives, technical solutions and advantages of the present invention more clear, the embodiments of the present invention will be further described in detail below with reference to the accompanying drawings.

[0051] A method for calculating and correcting the linear load distribution (normal force coefficient distribution density along the axial direction) of a launch vehicle is provided, which involves the calculation of the launch vehicle load. The specific steps are as follows:

[0052] (1) The CFD method is used to perform numerical simulation on the aerodynamic characteristics of the launch vehicle three-dimensional model. The aerodynamic simulation coordinate system is as follows: Figure 1As shown in the figure, based on the numerical simulation results, the aerodynamic characteristic parameters of the rocket body surface are output, including the axial coordinate X of the center of each grid unit on the rocket body surface. j 、Pressure P j , normal component of shear stress τ jy , unit area A j And the unit normal projection area A jy .

[0053] (2) The aerodynamic line load distribution is calculated using the piecewise integration method. The specific steps are as follows:

[0054] (a) Set the aerodynamic line distribution load deployment node sequence x along the axial direction of the rocket body i , i=1,2,...,N. Distribution and deployment node sequence x i The nodes should be appropriately encrypted in areas with drastic flow changes, and appropriately sparse in areas with gentle flow changes. For areas with drastic flow changes such as the nose cone of the satellite fairing, the transition area of ​​the cone column, the inverted cone, and the tail installation area, the deployment nodes need to be encrypted. The distance between the two deployment nodes after encryption should be no less than 3 times the characteristic size of the local grid unit; for areas with gentle flow changes such as the first-stage barrel section, the deployment nodes can be appropriately sparse. The distance between the two deployment nodes after sparseness should be no more than 20 times the characteristic size of the local grid unit.

[0055] (b) According to the deployment node sequence x i Set the rocket body axial segment control volume Δx i , the schematic diagram of the axial segmented control body is as follows Figure 2 As shown, each segment control volume Δx is calculated at the same time i The center coordinates s i :

[0056]

[0057]

[0058] (c) Axial segment control volume Δx of the rocket body i All CFD surface mesh elements within are integrated, and the segment normal force coefficient ΔC is calculated according to the following formula ni , where Q ref is the reference dynamic pressure, A ref is the reference area of ​​the rocket, and the schematic diagram of the integral of the surface pressure and shear stress of the rocket body is as follows Figure 3 As shown:

[0059]

[0060] (d) In the axial segment control volume Δx i The linear load distribution (normal force coefficient along the axial distribution density) C' is calculated according to the following formula ni :

[0061]

[0062] (3) According to the target normal force coefficient C n,t (The normal force coefficient is the integral result of the normal force coefficient distribution density) and the target pressure center position X p,t Correct the load distribution of the aerodynamic line. The specific steps are as follows:

[0063] (a) The normal force coefficient C of the rocket body can be calculated by the aerodynamic line load distribution n and the pressure center position X p :

[0064] C n =∑C′ ni Δx i

[0065] C n X p =∑C′ ni Δx i ·s i

[0066] (b) The target normal force coefficient C of the rocket body is obtained by numerical simulation, wind tunnel test or deviation design. n,t and target pressure center position X p,t , then there is a target aerodynamic line load distribution C′ ni,t satisfy:

[0067] C n,t =∑C′ ni,t Δx i

[0068] C n,t X p,t =∑C′ ni,t Δx i ·s i

[0069] (c) The corrected target aerodynamic line load distribution C′ can be calculated by the following formula: ni,t :

[0070] C′ ni,t =[k(xX p )+b]C′ ni

[0071] Where:

[0072]

[0073]

[0074] (4) Output the calculation results of the aerodynamic linear load distribution, including the center coordinates s of each segment control body in the axial direction of the rocket body i And the normal force coefficient along the axial distribution density value C' ni,t .

[0075] The following is further described in conjunction with specific embodiments:

[0076] S1. Use CFD method to perform numerical simulation on the aerodynamic characteristics of a three-dimensional model of a launch vehicle. Based on the numerical simulation results, output the aerodynamic characteristic parameters of the rocket body surface, including the axial coordinate X of the center of each grid unit on the rocket body surface. j 、Pressure P j , normal component of shear stress τ jy , unit area A j And the unit normal projection area A jy And other information.

[0077] S2, set 69 aerodynamic line distribution load deployment nodes along the axial direction of the rocket body, and set 69 axial segment control bodies Δx according to the deployment nodes i , and calculate each segmented control volume Δx i The center coordinates s i ; Axial segment control volume Δx for the rocket body i All CFD surface mesh elements within the segment are integrated to obtain the segment normal force coefficient, and then the axial distribution density of the normal force coefficient is obtained.

[0078] S3. Set the target normal force coefficient and target pressure center position according to the wind tunnel test results, and correct the aerodynamic line load distribution to obtain the corrected target aerodynamic line load distribution C′ ni,t .

[0079] S4. Output the calculation results of the aerodynamic load distribution, including the center coordinates s of each segment control body in the axial direction of the rocket body i And the normal force coefficient along the axial distribution density value C' ni,t Comparison of aerodynamic load distribution of a certain working condition calculated by the new method and the existing technology Figure 4 As shown in the figure, it can be seen that the new method provided by the present invention can better capture the aerodynamic characteristics at the place where the shape changes dramatically, and obtain a more accurate aerodynamic line load distribution.

[0080] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

[0081] 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.

Claims

1. A method for calculating and correcting the linear load distribution of a launch vehicle, characterized in that: include: (1) Use CFD method to perform numerical simulation on the aerodynamic characteristics of the launch vehicle three-dimensional model, and output the aerodynamic characteristic parameters of the rocket body surface based on the numerical simulation results; (2) Calculate the aerodynamic line load distribution using the piecewise integration method; The calculation methods for aerodynamic line load distribution include: (a) Set the aerodynamic line distribution load deployment node sequence x along the axial direction of the rocket body i , i=1,2,…,N; (b) According to the deployment node sequence x i Set the rocket body axial segment control volume Δx i , and calculate each segmented control volume Δx i The center coordinates s i : (c) Axial segment control volume Δx of the rocket body i All CFD surface mesh elements within are integrated, and the segment normal force coefficient ΔC is calculated according to the following formula ni , where Q ref is the reference dynamic pressure, A ref The reference area of ​​the rocket is: (d) In the axial segment control volume Δx i The normal force coefficient along the axial distribution density C is calculated according to the following formula: n ' i , where C n is the normal force coefficient: (3) Correcting the aerodynamic line load distribution according to the target normal force coefficient and the target pressure center position; The aerodynamic line load distribution correction methods include: (a) The normal force coefficient C of the rocket body can be calculated by the aerodynamic line load distribution n and the pressure center position X p : C n =∑C n ′ i ·Δx i C n X p =ΣC n ′ i ·Δx i ·s i (b) Obtain the target normal force coefficient C of the rocket body through numerical simulation, wind tunnel test or deviation design n,t and target pressure center position X p,t , then there is a target aerodynamic line load distribution C n ' i,t satisfy: C n,t =∑C n ′ i,t ·Δx i C n,t X p,t =ΣC n ′ i,t ·Δx i ·s i (c) The corrected target aerodynamic line load distribution C can be calculated by the following formula: n ' i,t : C n ′ i,t =[k(x-X p )+b]C n ′ i Where: (4) Output the calculation results of aerodynamic line load distribution.

2. The method according to claim 1, characterized in that: The aerodynamic characteristic parameters of the rocket body surface include at least the axial coordinate X of the center of each grid unit on the rocket body surface. j 、Pressure P j , normal component of shear stress τ jy , unit area A j And the unit normal projection area A jy .

3. The method according to claim 1, characterized in that Distribute and deploy node sequence x i The density is increased in areas with drastic flow changes and decreased in areas with gentle flow changes.

4. The method according to claim 3, characterized in that For the satellite fairing nose cone, cone-column transition area, inverted cone, and tail wing installation area, the deployment nodes are encrypted, and the distance between two deployment nodes after encryption is not less than 3 times the local grid unit characteristic size; for the first-stage barrel section area, the nodes are sparsely deployed, and the distance between two deployment nodes after sparseness should not be greater than 20 times the local grid unit characteristic size.

5. The method according to claim 1, characterized in that: The calculation results of aerodynamic linear load distribution include at least the center coordinates s of each segment control body in the axial direction of the rocket body i And the normal force coefficient along the axial distribution density C n ' i,t .

6. The method according to any one of claims 1 to 5, characterized in that The linear load distribution of a launch vehicle includes at least the axial distribution density of the normal force coefficient.

7. A computer-readable storage medium having computer program instructions stored thereon, wherein when the computer program instructions are loaded and executed by a processor, the processor is caused to execute the method according to any one of claims 1 to 5.

8. An electronic device comprising: processor; as well as a memory for storing computer program instructions; When the computer program instructions are loaded and executed by the processor, the processor performs the method according to any one of claims 1 to 5.

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