CN111985169A - Method for modeling aerodynamic characteristics of near-axisymmetric aircraft - Google Patents

Method for modeling aerodynamic characteristics of near-axisymmetric aircraft Download PDF

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CN111985169A
CN111985169A CN202010871303.6A CN202010871303A CN111985169A CN 111985169 A CN111985169 A CN 111985169A CN 202010871303 A CN202010871303 A CN 202010871303A CN 111985169 A CN111985169 A CN 111985169A
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马玉海
吴炜平
廉洁
樊鹏飞
史晓宁
杨毅强
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Beijing Zhongke Aerospace Technology Co Ltd
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Abstract

The application relates to the technical field of aircraft research and development, in particular to a method for modeling aerodynamic characteristics of a paraxial symmetry aircraft, which comprises the following steps: calculating the aerodynamic characteristics of the paraxial symmetry aircraft body; carrying out continuous rolling wind tunnel test on the paraxial symmetry aircraft body and the protrusion attached to the aircraft body to obtain the aerodynamic characteristics of the protrusion; calculating the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface of the paraxial symmetry aircraft; and superposing the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface and the aerodynamic characteristics of the protrusion on the aerodynamic characteristics of the aircraft body so as to realize the near-axis symmetry aircraft aerodynamic characteristic modeling. The method and the device can reduce the calculation amount of the pneumatic characteristic modeling of the paraxial symmetry aircraft, thereby quickly obtaining the aerodynamic force and the aerodynamic moment and realizing the quick modeling of the pneumatic characteristic.

Description

Method for modeling aerodynamic characteristics of near-axisymmetric aircraft
Technical Field
The application relates to the technical field of aircraft research and development, in particular to a method for modeling aerodynamic characteristics of a paraxial symmetry aircraft.
Background
The carrier rocket or missile is usually based on a multi-section conical rotating body, bulges such as a separation device, a pipeline, a cable protection cover and the like are added on the surface according to the structural layout requirement, and pneumatic parts such as a pneumatic stabilizing surface or a control surface and the like are added according to the control requirement, so that the carrier rocket or missile is a paraxial symmetry aircraft.
In the early stages of development of aircraft, it was necessary to give approximate aerodynamic characteristics (aerodynamic force, aerodynamic moment) for preliminary performance evaluation. In order to obtain data capable of describing the aerodynamic characteristics of the paraxial symmetry aircraft and establish a mathematical model of the aerodynamic force and the aerodynamic moment applied to the aircraft, a wind tunnel test is required to simulate the influence of the protrusions and the aerodynamic components.
Aerodynamic force and aerodynamic moment under a plurality of aerodynamic characteristic data are obtained through wind tunnel tests, dimensionless aerodynamic force coefficient and aerodynamic moment coefficient under the aerodynamic characteristic data are obtained according to the aerodynamic force and the aerodynamic moment, and accordingly a data table corresponding to the aerodynamic force coefficient, the aerodynamic moment coefficient and the aerodynamic characteristic data is built.
I.e. in the form of a number of aerodynamic characteristic data (typical mach number Ma, angle of attack alpha, angle of sideslip beta and deflection angle of the individual aerodynamic componentiAnd i is 1, 2, …, N (such as deflection angle of an air rudder and a grid rudder, which is called rudder deflection angle for short)) is taken as an independent variable, and a dimensionless aerodynamic coefficient (axial force coefficient C) is givenACoefficient of normal force CNCoefficient of lateral force CZ) Aerodynamic moment coefficient (rolling moment coefficient m)XCoefficient of yawing moment mYCoefficient of pitching moment mZ) The data table of (1).
According to aerodynamic characteristic data (Mach number Ma, angle of attack alpha, angle of sideslip beta, angle of deflection)i) In the pneumatic powerThe aerodynamic coefficient and the aerodynamic moment coefficient under the aerodynamic characteristic data are obtained by interpolation in the force coefficient and aerodynamic moment coefficient data table, and then the formula is followed
Figure BDA0002651181410000021
Corresponding aerodynamic force and aerodynamic moment are obtained.
Wherein the dynamic pressure
Figure BDA0002651181410000022
ρ is the atmospheric density, SMFor reference area,/KFor reference length, FxComponent of aerodynamic force along the coordinate axis x, FyComponent of aerodynamic force along the coordinate axis y, FzComponent of aerodynamic force along the coordinate axis z, MxComponent of aerodynamic moment along the coordinate axis x, MyComponent of aerodynamic moment along the coordinate axis y, MzIs the component of the aerodynamic moment along the coordinate axis z.
However, because of the asymmetric factors, the state combinations of the independent variables are many, in order to obtain a data table of the aerodynamic coefficient and the aerodynamic moment coefficient, the required calculation state also increases exponentially with the number of the independent variables, which causes a great increase in calculation amount, resulting in a long working period and high cost, and for the initial design, accurate modeling data is not required, only the main influencing factors need to be evaluated, and the variation probability is high, and a great waste is caused by repeatedly obtaining the aerodynamic characteristics.
Therefore, how to reduce the calculation amount of the paraxial symmetry aircraft aerodynamic characteristic modeling so as to quickly obtain aerodynamic force and aerodynamic moment is a technical problem to be solved by the technical personnel in the field at present.
Disclosure of Invention
The application provides a method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft, which is used for reducing the calculation amount of the modeling of the aerodynamic characteristics of the near-axisymmetric aircraft, so that aerodynamic force and aerodynamic moment are quickly obtained, and the quick modeling of the aerodynamic characteristics is realized.
In order to solve the technical problem, the application provides the following technical scheme:
a method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft comprises the following steps: calculating the aerodynamic characteristics of the paraxial symmetry aircraft body; carrying out continuous rolling wind tunnel test on the paraxial symmetry aircraft body and the protrusion attached to the aircraft body to obtain the aerodynamic characteristics of the protrusion; calculating the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface of the paraxial symmetry aircraft; and superposing the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface and the aerodynamic characteristics of the protrusion on the aerodynamic characteristics of the aircraft body so as to realize the near-axis symmetry aircraft aerodynamic characteristic modeling.
The method for modeling the aerodynamic characteristics of a paraxial symmetry aircraft as described above, wherein preferably the aerodynamic characteristics of the protrusion are obtained, comprises the following sub-steps: carrying out a continuous rolling wind tunnel test to obtain a total attack angle and a rolling angle in the whole rotating direction, and aerodynamic force and aerodynamic moment under the total attack angle and the rolling angle; converting the total attack angle and the roll angle into an attack angle and a sideslip angle; constructing a data table of aerodynamic coefficient and aerodynamic moment coefficient according to the obtained attack angle and sideslip angle, and corresponding aerodynamic force and aerodynamic moment; in the aerodynamic coefficient and aerodynamic moment coefficient data table, the aerodynamic coefficient and the aerodynamic moment coefficient under the corresponding attack angle and sideslip angle are obtained through interpolation, so that the aerodynamic characteristics of the protrusion under the attack angle and the sideslip angle are obtained.
The method for modeling the aerodynamic characteristics of the paraxial symmetry aircraft as described above, wherein the total attack angle and the roll angle are preferably converted into the attack angle and the sideslip angle, comprises the following sub-steps: respectively rotating the speed vector V of the airflow from an airflow coordinate system A along a y axis and a z axis by a sideslip angle beta and an attack angle alpha, and converting the speed vector V into an aircraft body coordinate system B; defining a roll angle phi and a total attack angle omega as rotation angles along an x axis and a z axis respectively, and converting a velocity vector V of the airflow from an airflow coordinate system A into an aircraft body coordinate system B by rotating the total attack angle omega and the roll angle phi along the z axis and the x axis respectively; applying element coordinate transformation matrixes around x, y and z coordinate axes to the process of transforming the velocity vector V of the airflow into an aircraft body coordinate system B to obtain the conversion relation between the total attack angle omega and the rolling angle phi and the attack angle alpha and the sideslip angle beta; and converting the total attack angle omega and the roll angle phi into the attack angle alpha and the sideslip angle beta according to conversion relational expressions of the total attack angle omega and the roll angle phi and the attack angle alpha and the sideslip angle beta.
The method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft as described above, wherein preferably the primitive coordinate transformation matrix around the x, y, z coordinate axes is
Figure BDA0002651181410000031
Figure BDA0002651181410000032
Wherein, Cx(θx) For rotation of theta along the x-axisxPrimitive coordinate transformation matrix of angle, Cy(θy) For rotation of theta along the y-coordinate axisyPrimitive coordinate transformation matrix of angle, Cz(θz) For rotation of theta along the z-coordinate axiszPrimitive coordinate transformation matrix of the angle.
The method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft as described above, wherein the velocity vector V of the airflow after being transformed into the aircraft body coordinate system B is preferably VB＝Cz(α)Cy(β)VA＝Cx(φ)Cz(Ω)VAIn which C isz(α) is a primitive coordinate transformation matrix rotated by an angle α along the z coordinate axis, Cy(beta) is a primitive coordinate transformation matrix rotated by an angle beta along the y coordinate axis, Cx(phi) is a primitive coordinate transformation matrix rotated by a phi angle along the x coordinate axis, Cz(omega) is a primitive coordinate transformation matrix rotated by an angle omega along the z coordinate axis, VAIs a coordinate vector of the velocity vector V of the air flow in the air flow coordinate system a.
The method for modeling the aerodynamic characteristics of the paraxial symmetry aircraft as described above, wherein the conversion relation between the total attack angle Ω and the roll angle Φ and the attack angle α and the sideslip angle β is preferably
Figure BDA0002651181410000041
Figure BDA0002651181410000042
The paraxial symmetry aircraft aerodynamic characteristics as described aboveMethod of sexual modelling wherein, preferably, the method is by formula
Figure BDA0002651181410000043
Calculating to obtain the aerodynamic characteristics of the protrusion, wherein alpha is an attack angle, beta is a sideslip angle, Ma is a Mach number, and CAIs the axial force coefficient, CNIs the normal force coefficient, CZIs the coefficient of lateral force, mXIs the rolling moment coefficient, mYIs the yaw moment coefficient, mZIs pitch moment coefficient, dynamic pressure
Figure BDA0002651181410000044
ρ is the atmospheric density, SMFor reference area,/KFor reference length, FxComponent of aerodynamic force along the coordinate axis x, FyComponent of aerodynamic force along the coordinate axis y, FzComponent of aerodynamic force along the coordinate axis z, MxComponent of aerodynamic moment along the coordinate axis x, MyComponent of aerodynamic moment along the coordinate axis y, MzIs the component of the aerodynamic moment along the coordinate axis z.
The method for modeling the aerodynamic characteristics of a paraxial symmetric aircraft as described above, wherein preferably the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface of the paraxial symmetric aircraft are calculated, comprises the following sub-steps: calculating a traction attack angle and a traction sideslip angle of the attitude of the pneumatic stabilizer and/or the control surface relative to the airflow, which is subjected to traction change along with the attitude of the aircraft body when the pneumatic stabilizer and/or the control surface is at a zero position; calculating according to the involvement attack angle and the involvement sideslip angle to obtain a local attack angle and a local sideslip angle; and inserting values of the aerodynamic coefficient and the aerodynamic moment coefficient under the local attack angle and the local sideslip angle in an aerodynamic coefficient and aerodynamic moment coefficient data table of the aerodynamic component according to the local attack angle and the local sideslip angle so as to obtain the aerodynamic characteristics of the aerodynamic component under the local attack angle and the local sideslip angle.
The method for modeling the aerodynamic characteristics of a paraxial symmetric aircraft as described above, wherein it is preferable to involve the angle of attack
Figure BDA0002651181410000051
Lead-in sideslip angle
Figure BDA0002651181410000052
Wherein alpha is an angle of attack, beta is a sideslip angle, phiiThe mounting azimuth angle of the aerodynamic component on the aircraft body.
The method for modeling the aerodynamic characteristics of a paraxial symmetric aircraft as described above, wherein it is preferred that the local angle of attack is such that if the aerodynamic component is a fixed stabilizer, the aerodynamic component has a local angle of attack
Figure BDA0002651181410000053
Local sideslip angle
Figure BDA0002651181410000054
If the pneumatic part is a control surface, its deflection angle isiCorrecting the local angle of attack to
Figure BDA0002651181410000055
Local sideslip angle
Figure BDA0002651181410000056
Compared with the background art, the method for modeling the aerodynamic characteristics of the paraxial symmetry aircraft separates the acquisition of the aerodynamic characteristics of the protrusion from the acquisition of the aerodynamic characteristics of the aerodynamic stabilizer and the control surface, so that the deflection angle of the aerodynamic part is reduced when an aerodynamic coefficient and aerodynamic moment coefficient data table of the protrusion is constructediThe parameter reduces the state combination of independent variables caused by asymmetric factors, reduces the calculation amount of the near-axisymmetric aircraft aerodynamic characteristic modeling, can further quickly obtain aerodynamic force and aerodynamic moment, and realizes the quick modeling of the near-axisymmetric aircraft aerodynamic characteristic.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments described in the present invention, and other drawings can be obtained by those skilled in the art according to the drawings.
FIG. 1 is a flow chart of a method for modeling the aerodynamic characteristics of a paraxial symmetric aircraft provided by an embodiment of the present application;
FIG. 2 is a flow chart for obtaining aerodynamic characteristics of a projection provided by an embodiment of the present application;
FIG. 3 is a flow chart for converting a total angle of attack, roll angle to an angle of attack, sideslip angle provided by an embodiment of the present application;
FIG. 4 is a flow chart for calculating aerodynamic characteristics of an aerodynamic stabilizer and/or control surface provided by an embodiment of the present application.
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings, wherein like or similar reference numerals refer to the same or similar elements or elements having the same or similar function throughout. The embodiments described below with reference to the drawings are illustrative only and should not be construed as limiting the invention.
Referring to fig. 1, fig. 1 is a flowchart of a method for modeling an aerodynamic characteristic of a paraxial symmetry aircraft according to an embodiment of the present application.
The application provides a method for modeling aerodynamic characteristics of a near-axisymmetric aircraft, which comprises the following steps:
step S110, calculating the aerodynamic characteristics of the paraxial symmetry aircraft body;
the multi-section conical rotating body of the carrier rocket is a main body part of the carrier rocket, and can be calculated according to a traditional engineering algorithm or a small-scale CFD (Computational Fluid Dynamics) model to obtain the aerodynamic characteristics of the near-axisymmetric aircraft body, namely obtain the aerodynamic force and the aerodynamic moment of the near-axisymmetric aircraft body.
S120, carrying out a continuous rolling wind tunnel test on the paraxial symmetry aircraft body and the protrusion attached to the aircraft body to obtain the aerodynamic characteristics of the protrusion;
referring to fig. 2, fig. 2 is a flow chart for acquiring aerodynamic characteristics of a protrusion according to an embodiment of the present disclosure;
specifically, the method for obtaining the aerodynamic characteristics of the protrusion by carrying out a continuous rolling wind tunnel test comprises the following substeps:
step S121, carrying out a continuous rolling wind tunnel test, and acquiring a total attack angle and a rolling angle in the whole rotating direction, and aerodynamic force and aerodynamic moment under the total attack angle and the rolling angle;
for the protrusion, due to the fact that the influence mechanism is complex, the CFD calculation scale requirement is high, the time is long, wind tunnel tests generally need to be carried out, however, due to the fact that the protrusion mainly influences the rolling moment coefficient, and the method is used for preliminary performance evaluation of aerodynamic characteristics in the early development stage of an aircraft, the influence of the protrusion on other parameters is ignored, continuous rolling wind tunnel tests are carried out, and therefore the total attack angle omega and the rolling angle phi in the whole rotating direction and aerodynamic force and aerodynamic moment measured under different total attack angles omega and different rolling angles phi can be obtained through one wind tunnel test, and the times of the wind tunnel tests are further reduced.
In addition, the wind tunnel test is carried out according to the states of the total attack angle omega and the roll angle phi, which is beneficial to automatic operation, but in order to carry out complete simulation on various factors in the aircraft dynamic model, the description forms of the total attack angle omega and the roll angle phi are preferably unified into the description forms of the attack angle alpha and the sideslip angle beta.
Step S122, converting the total attack angle and the roll angle into an attack angle and a sideslip angle;
for the same space three-dimensional vector, different coordinates are required to be represented among different coordinate systems, and coordinate values among the different coordinate systems can be converted according to a coordinate transformation matrix. Specifically, the primitive coordinate transformation matrices around each coordinate axis x, y, and z are respectively:
Figure BDA0002651181410000071
Figure BDA0002651181410000072
Figure BDA0002651181410000073
wherein, Cx(θx) For rotation of theta along the x-axisxPrimitive coordinate transformation matrix of angle, Cy(θy) For rotation of theta along the y-coordinate axisyPrimitive coordinate transformation matrix of angle, Cz(θz) For rotation of theta along the z-coordinate axiszPrimitive coordinate transformation matrix of the angle.
For any coordinate transformation, it can be expressed as a composition of three independent successive transformations, e.g. by rotating the coordinate system X by θ along the three coordinate axes z → y → Xz，θyAnd thetaxThree angles, making them coincide with the coordinate system Y, and obtaining a synthetic coordinate transformation matrix of
Figure BDA0002651181410000074
In addition, a coordinate vector of a certain space vector r in the coordinate system X is represented as rXThe coordinate vector in the coordinate system Y is denoted as rY，rXAnd rYThe relationship of (1) is: wherein
Figure BDA0002651181410000081
Representing the coordinate component of the vector r along the coordinate axis p in the coordinate system q, then
Figure BDA0002651181410000082
Is the coordinate component of the vector r along the coordinate axis x in the coordinate system Y,
Figure BDA0002651181410000083
is the coordinate component of the vector r along the coordinate axis Y in the coordinate system Y,
Figure BDA0002651181410000084
is the coordinate component of the vector r along the coordinate axis z in the coordinate system Y,
Figure BDA0002651181410000085
is the coordinate component of the vector r along the coordinate axis X in the coordinate system X,
Figure BDA0002651181410000086
is the coordinate component of the vector r along the coordinate axis y in the coordinate system X,
Figure BDA0002651181410000087
is the coordinate component of the vector r along the coordinate axis z in the coordinate system X.
Referring to fig. 3, fig. 3 is a flowchart illustrating a process of converting a total angle of attack and a roll angle into an angle of attack and a sideslip angle according to an embodiment of the present application;
specifically, the method for converting the total attack angle and the roll angle into the attack angle and the sideslip angle comprises the following substeps:
step S1221, the velocity vector V of the air flow is determined from the air flow coordinate system
Figure BDA0002651181410000088
Respectively rotating a sideslip angle beta and an attack angle alpha along the y axis and the z axis, and transforming the sideslip angle beta and the attack angle alpha into an aircraft body coordinate system B, namely VB＝Cz(α)Cy(β)VA；
Specifically, the aircraft body coordinate system is represented as B, the airflow coordinate system is represented as a, and since the y-axes of the two coordinate systems (the aircraft body coordinate system B and the airflow coordinate system a) are in the same plane, only two successive rotations are needed for transformation, and the coordinate transformation matrix is
Figure BDA0002651181410000089
Alpha is angle of attack, beta is angle of sideslip, Cz(α) is a primitive coordinate transformation matrix rotated by an angle α along the z coordinate axis, CyAnd (beta) is a primitive coordinate transformation matrix which rotates by an angle beta along the y coordinate axis.
Then for the velocity vector V of the air flow the coordinate vector in the aircraft body coordinate system B is
Figure BDA00026511814100000810
Wherein
Figure BDA00026511814100000811
Is the coordinate component of the velocity vector V of the air flow along the coordinate axis x in the coordinate system B,
Figure BDA00026511814100000812
is the coordinate component of the velocity vector V of the air flow along the coordinate axis y in the coordinate system B,
Figure BDA00026511814100000813
is the coordinate component of the velocity vector V of the air flow along the coordinate axis z in the coordinate system B.
The coordinate vector of the velocity vector V of the air flow in the air flow coordinate system A is
Figure BDA0002651181410000091
Wherein the content of the first and second substances,
Figure BDA0002651181410000092
is the coordinate component of the velocity vector V of the air flow along the coordinate axis x in the coordinate system a,
Figure BDA0002651181410000093
is the coordinate component of the velocity vector V of the air flow along the coordinate axis y in the coordinate system a,
Figure BDA0002651181410000094
is the coordinate component of the velocity vector V of the air flow along the coordinate axis z in the coordinate system a.
Will VAConversion to VBI.e. by
Figure BDA0002651181410000095
Step S1222, defining a roll angle phi and a total attack angle omega as rotation angles along the x-axis and the z-axis, respectively, rotating the velocity vector V of the airflow from the airflow coordinate system A along the z-axis and the x-axis to the total attack angle omega and the roll angle phi, respectively, and transforming the velocity vector V into an aircraft body coordinate system B, namely VB＝Cx(φ)Cz(Ω)VA；
That is, VB＝Cz(α)Cy(β)VA＝Cx(φ)Cz(Ω)VA，Cx(phi) is a primitive coordinate transformation matrix rotated by a phi angle along the x coordinate axis, CzAnd (omega) is a primitive coordinate transformation matrix which rotates by an omega angle along the z coordinate axis.
S1223, applying the element coordinate transformation matrix around the coordinate axes x, y and z to the process of transforming the speed vector V of the airflow into the aircraft body coordinate system B to obtain the conversion relation between the total attack angle omega and the roll angle phi and the attack angle alpha and the sideslip angle beta;
will be provided with
Figure BDA0002651181410000096
Figure BDA0002651181410000097
Applied to the above formula VB＝Cz(α)Cy(β)VA＝Cx(φ)Cz(Ω)VAThereby obtaining a conversion relation between different angles defined by two transformation methods:
Figure BDA0002651181410000098
Figure BDA0002651181410000099
step S1224, converting the total angle of attack Ω and the roll angle Φ into an angle of attack α and a sideslip angle β according to the conversion relationship between the total angle of attack Ω and the roll angle Φ and the angle of attack α and the sideslip angle β.
Conversion relation formula of total attack angle omega, rolling angle phi, attack angle alpha and sideslip angle beta is utilized
Figure BDA0002651181410000101
Figure BDA0002651181410000102
Given any pair of angle definitions, such as angle of attack α and sideslip angle β, the equivalent total angle of attack Ω and roll angle φ can be found, and vice versa.
S123, constructing a data table of aerodynamic coefficient and aerodynamic moment coefficient according to the obtained attack angle and sideslip angle, and corresponding aerodynamic force and aerodynamic moment;
a plurality of total attack angles omega and rolling angles phi and a plurality of corresponding aerodynamic forces and aerodynamic moments are measured through a wind tunnel test, the total attack angles omega and the rolling angles phi are converted into attack angles alpha and sideslip angles beta, then a plurality of attack angles alpha and sideslip angles beta and a plurality of corresponding aerodynamic forces and aerodynamic moments can be obtained, and the formula is used for obtaining the aerodynamic forces and the aerodynamic moments
Figure BDA0002651181410000103
Aerodynamic coefficient and aerodynamic moment coefficient under corresponding attack angle alpha and sideslip angle beta can be obtained, and corresponding aerodynamic coefficient and aerodynamic moment coefficient data tables are constructed by taking the attack angle alpha and the sideslip angle beta as independent variables, wherein Mach number Ma and aerodynamic coefficient (axial force coefficient C)ACoefficient of normal force CNCoefficient of lateral force CZ) Coefficient of aerodynamic moment (coefficient of rolling moment m)XCoefficient of yawing moment mYCoefficient of pitching moment mZ) Dynamic pressure
Figure BDA0002651181410000104
ρ is the atmospheric density, SMFor reference area,/KFor reference length, FxComponent of aerodynamic force along the coordinate axis x, FyComponent of aerodynamic force along the coordinate axis y, FzComponent of aerodynamic force along the coordinate axis z, MxComponent of aerodynamic moment along the coordinate axis x, MyComponent of aerodynamic moment along the coordinate axis y, MzIs the component of the aerodynamic moment along the coordinate axis z.
And S124, in the aerodynamic coefficient and aerodynamic moment coefficient data tables, the aerodynamic coefficient and the aerodynamic moment coefficient under the corresponding attack angle and sideslip angle are obtained through interpolation, so that the aerodynamic characteristics of the protrusion under the attack angle and the sideslip angle are obtained.
Because the quantity of data measured by a wind tunnel test is limited, the data quantity in the constructed data table of the aerodynamic coefficient and the aerodynamic moment coefficient is also limited, and therefore, in order to obtain the aerodynamic coefficient and the aerodynamic moment coefficient under any attack angle and sideslip angle, the aerodynamic coefficient and the aerodynamic moment coefficient need to be acquiredInquiring aerodynamic coefficient and aerodynamic moment coefficient under corresponding attack angle and sideslip angle from aerodynamic coefficient and aerodynamic moment coefficient data table by interpolation method, and then passing through formula
Figure BDA0002651181410000111
The aerodynamic force and the aerodynamic moment of the protrusion, that is, the aerodynamic characteristics of the protrusion, are obtained through calculation.
S130, calculating the aerodynamic characteristics of an aerodynamic stabilizer and/or control surface of the paraxial symmetry aircraft;
referring to FIG. 4, FIG. 4 is a flow chart for calculating aerodynamic characteristics of an aerodynamic stabilizer and/or control surface according to an embodiment of the present disclosure;
specifically, the method for calculating the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface of the paraxial symmetry aircraft comprises the following sub-steps:
s131, calculating a involvement attack angle and a sideslip angle of the aerodynamic stabilizer and/or the control surface, which are involved and changed along with the attitude of the aircraft body relative to the attitude of the aerodynamic stabilizer and/or the control surface when the aerodynamic stabilizer and/or the control surface is at a zero position;
specifically, the local coordinate system of a certain pneumatic component i is represented as RiThe air flow coordinate system is expressed as A, the velocity vector of a certain pneumatic component i relative to the air flow is W, and the velocity vector W of a certain pneumatic component i relative to the air flow is in the local coordinate system RiThe coordinate vector of (1) is
Figure BDA0002651181410000112
Wherein the content of the first and second substances,
Figure BDA0002651181410000113
is a velocity vector W in a coordinate system RiOf the coordinate component along the coordinate axis x,
Figure BDA0002651181410000114
is a velocity vector W in a coordinate system RiThe coordinate component of (a) along the coordinate axis y,
Figure BDA0002651181410000115
is a velocity vector W in a coordinate system RiMiddle coordinate division along coordinate axis zAmount of the compound (A).
The coordinate vector of a certain pneumatic component i in the airflow coordinate system A relative to the velocity vector W of the airflow is
Figure BDA0002651181410000116
Wherein the content of the first and second substances,
Figure BDA0002651181410000117
is the coordinate component of the velocity vector W along the coordinate axis x in the coordinate system a,
Figure BDA0002651181410000118
is the coordinate component of the velocity vector W along the coordinate axis y in the coordinate system a,
Figure BDA0002651181410000119
is the coordinate component of the velocity vector W along the coordinate axis z in the coordinate system a.
And the number of the first and second electrodes,
Figure BDA0002651181410000121
wherein phi isiFor the installation azimuth of the aerodynamic component on the aircraft body,
Figure BDA0002651181410000122
for the angle of attack involved for the pneumatic component,
Figure BDA0002651181410000123
the angle of the trailing sideslip of the pneumatic component.
Transforming the coordinates of the elements around the x, y and z coordinate axes into a matrix
Figure BDA0002651181410000124
Figure BDA0002651181410000125
Is applied to
Figure BDA0002651181410000126
In (1) obtaining
Figure BDA0002651181410000127
Step S132, calculating according to the involvement attack angle and the involvement sideslip angle to obtain a local attack angle and a local sideslip angle;
in particular, if the aerodynamic element is a fixed stabilizer, its local angle of attack
Figure BDA0002651181410000128
Local sideslip angle
Figure BDA0002651181410000129
If the pneumatic part is a movable control surface, its deflection angle isiCorrecting the local angle of attack to
Figure BDA00026511814100001210
Local sideslip angle
Figure BDA00026511814100001211
Step S133, according to the local attack angle and the local sideslip angle, the aerodynamic coefficient and the aerodynamic moment coefficient under the local attack angle and the local sideslip angle are obtained by inserting values in the aerodynamic coefficient and the aerodynamic moment coefficient data table of the pneumatic component, so as to obtain the aerodynamic characteristics of the pneumatic component under the local attack angle and the local sideslip angle.
According to the above local angle of attack alphaiLocal sideslip angle betaiIn the aerodynamic coefficient and aerodynamic moment coefficient data table of the aerodynamic component, approximate aerodynamic coefficient and aerodynamic moment coefficient are calculated through interpolation.
Then, by the formula
Figure BDA00026511814100001212
Calculating to obtain the aerodynamic force and the aerodynamic moment of the pneumatic component, namely the aerodynamic characteristics of the pneumatic component, wherein the dimensionless aerodynamic coefficient (comprising the axial force coefficient C) of the pneumatic component iiACoefficient of normal force CiNCoefficient of lateral force CiZ) Aerodynamic moment coefficient (including rolling moment coefficient m)iXCoefficient of yawing moment miYCoefficient of pitching moment miZ) Root of Chinese ginsengConsideration (reference length l)iKReference area SiM)，FixComponent of aerodynamic force along the coordinate axis x, FiyComponent of aerodynamic force along the coordinate axis y, FizComponent of aerodynamic force along the coordinate axis z, MixComponent of aerodynamic moment along the coordinate axis x, MiyComponent of aerodynamic moment along the coordinate axis y, MizIs the component of the aerodynamic moment along the coordinate axis z.
And S140, superposing the aerodynamic characteristics of the protrusions and the aerodynamic characteristics of the aerodynamic stabilizer or the control surface on the aerodynamic characteristics of the aircraft body to realize the modeling of the aerodynamic characteristics of the near-axis symmetric aircraft.
The aerodynamic characteristics of the aerodynamic component under the local attack angle and the local sideslip angle are converted into a matrix through the coordinate of the aerodynamic component to the aircraft body
Figure BDA0002651181410000131
The aerodynamic characteristics of the aerodynamic component at the attack angle alpha and the sideslip angle beta are converted.
Alternatively, the influence of the reference amount of the aircraft may be newly dimensionless in advance according to the reference amount of the aircraft, and the influence may be corrected to the aerodynamic coefficient of the entire aircraft.
Since the aerodynamic characteristics of the protrusions are obtained separately from the aerodynamic characteristics of the aerodynamic stabilizer and control surface, i.e. the aerodynamic characteristics of the protrusions are obtained separately, the deflection angle of the aerodynamic part is reduced when constructing the data sheet of aerodynamic coefficient and aerodynamic moment coefficientiThe parameter reduces the state combination of independent variables caused by asymmetric factors, reduces the calculation amount of the near-axisymmetric aircraft aerodynamic characteristic modeling, can further quickly obtain aerodynamic force and aerodynamic moment, and realizes the quick modeling of the near-axisymmetric aircraft aerodynamic characteristic.
It will be evident to those skilled in the art that the invention is not limited to the details of the foregoing illustrative embodiments, and that the present invention may be embodied in other specific forms without departing from the spirit or essential attributes thereof. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive, the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein. Any reference sign in a claim should not be construed as limiting the claim concerned.
Furthermore, it should be understood that although the present description refers to embodiments, not every embodiment may contain only a single embodiment, and such description is for clarity only, and those skilled in the art should integrate the description, and the embodiments may be combined as appropriate to form other embodiments understood by those skilled in the art.

Claims (10)

1. A method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft is characterized by comprising the following steps:
calculating the aerodynamic characteristics of the paraxial symmetry aircraft body;
carrying out continuous rolling wind tunnel test on the paraxial symmetry aircraft body and the protrusion attached to the aircraft body to obtain the aerodynamic characteristics of the protrusion;
calculating the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface of the paraxial symmetry aircraft;
and superposing the aerodynamic characteristics of the aerodynamic stabilizer and/or control surface and the aerodynamic characteristics of the protrusion on the aerodynamic characteristics of the aircraft body so as to realize the near-axis symmetry aircraft aerodynamic characteristic modeling.
2. The method of modeling the aerodynamic properties of a paraxial symmetric aircraft of claim 1, wherein obtaining the aerodynamic properties of the protrusion comprises the substeps of:
carrying out a continuous rolling wind tunnel test to obtain a total attack angle and a rolling angle in the whole rotating direction, and aerodynamic force and aerodynamic moment under the total attack angle and the rolling angle;
converting the total attack angle and the roll angle into an attack angle and a sideslip angle;
constructing a data table of aerodynamic coefficient and aerodynamic moment coefficient according to the obtained attack angle and sideslip angle, and corresponding aerodynamic force and aerodynamic moment;
in the aerodynamic coefficient and aerodynamic moment coefficient data table, the aerodynamic coefficient and the aerodynamic moment coefficient under the corresponding attack angle and sideslip angle are obtained through interpolation, so that the aerodynamic characteristics of the protrusion under the attack angle and the sideslip angle are obtained.
3. The method of modeling near-axisymmetric aircraft aerodynamic properties of claim 2, wherein converting the total angle of attack, roll angle to angle of attack, sideslip angle comprises the substeps of:
respectively rotating the speed vector V of the airflow from an airflow coordinate system A along a y axis and a z axis by a sideslip angle beta and an attack angle alpha, and converting the speed vector V into an aircraft body coordinate system B;
defining a roll angle phi and a total attack angle omega as rotation angles along an x axis and a z axis respectively, and converting a velocity vector V of the airflow from an airflow coordinate system A into an aircraft body coordinate system B by rotating the total attack angle omega and the roll angle phi along the z axis and the x axis respectively;
applying element coordinate transformation matrixes around x, y and z coordinate axes to the process of transforming the velocity vector V of the airflow into an aircraft body coordinate system B to obtain the conversion relation between the total attack angle omega and the rolling angle phi and the attack angle alpha and the sideslip angle beta;
and converting the total attack angle omega and the roll angle phi into the attack angle alpha and the sideslip angle beta according to conversion relational expressions of the total attack angle omega and the roll angle phi and the attack angle alpha and the sideslip angle beta.
4. The method of claim 3, wherein the primitive coordinate transformation matrix around the x, y, z coordinate axes is
Figure FDA0002651181400000021
Figure FDA0002651181400000022
Wherein, Cx(θx) For rotation of theta along the x-axisxPrimitive coordinate transformation moment of angleArray, Cy(θy) For rotation of theta along the y-coordinate axisyPrimitive coordinate transformation matrix of angle, Cz(θz) For rotation of theta along the z-coordinate axiszPrimitive coordinate transformation matrix of the angle.
5. The method for modeling the aerodynamic characteristics of a near-axisymmetric aircraft according to claim 3 or 4, wherein the velocity vector V of the airflow after transformation into the aircraft body coordinate system B is VB＝Cz(α)Cy(β)VA＝Cx(φ)Cz(Ω)VAIn which C isz(α) is a primitive coordinate transformation matrix rotated by an angle α along the z coordinate axis, Cy(beta) is a primitive coordinate transformation matrix rotated by an angle beta along the y coordinate axis, Cx(phi) is a primitive coordinate transformation matrix rotated by a phi angle along the x coordinate axis, Cz(omega) is a primitive coordinate transformation matrix rotated by an angle omega along the z coordinate axis, VAIs a coordinate vector of the velocity vector V of the air flow in the air flow coordinate system a.
6. The method for modeling aerodynamic characteristics of a paraxial symmetry aircraft according to claim 3 or 4, wherein the conversion relation between the total attack angle Ω and the roll angle Φ and the attack angle α and the sideslip angle β is
Figure FDA0002651181400000023
7. The method for modeling the aerodynamic properties of a near-axisymmetric aircraft of any of claims 2-4, characterized by the formula
Figure FDA0002651181400000031
Calculating to obtain the aerodynamic characteristics of the protrusion, wherein alpha is an attack angle, beta is a sideslip angle, Ma is a Mach number, and CAIs the axial force coefficient, CNIs the normal force coefficient, CZIs the coefficient of lateral force, mXIs the rolling moment coefficient, mYIs the yaw moment coefficient, mZIs pitch moment coefficient, dynamic pressure
Figure FDA0002651181400000032
ρ is the atmospheric density, SMFor reference area,/KFor reference length, FxComponent of aerodynamic force along the coordinate axis x, FyComponent of aerodynamic force along the coordinate axis y, FzComponent of aerodynamic force along the coordinate axis z, MxComponent of aerodynamic moment along the coordinate axis x, MyComponent of aerodynamic moment along the coordinate axis y, MzIs the component of the aerodynamic moment along the coordinate axis z.
8. The method for modeling the aerodynamic characteristics of a paraxial symmetric aircraft according to any one of claims 1 to 4, wherein calculating the aerodynamic characteristics of the aerodynamic stabilizer and/or control surfaces of a paraxial symmetric aircraft comprises the substeps of:
calculating a traction attack angle and a traction sideslip angle of the attitude of the pneumatic stabilizer and/or the control surface relative to the airflow, which is subjected to traction change along with the attitude of the aircraft body when the pneumatic stabilizer and/or the control surface is at a zero position;
calculating according to the involvement attack angle and the involvement sideslip angle to obtain a local attack angle and a local sideslip angle;
and inserting values of the aerodynamic coefficient and the aerodynamic moment coefficient under the local attack angle and the local sideslip angle in an aerodynamic coefficient and aerodynamic moment coefficient data table of the aerodynamic component according to the local attack angle and the local sideslip angle so as to obtain the aerodynamic characteristics of the aerodynamic component under the local attack angle and the local sideslip angle.
9. The method of modeling near-axisymmetric aircraft aerodynamic properties of claim 8, wherein the angle of attack is implicated
Figure FDA0002651181400000033
Lead-in sideslip angle
Figure FDA0002651181400000034
Wherein alpha is an angle of attack, beta is a sideslip angle, phiiThe mounting azimuth angle of the aerodynamic component on the aircraft body.
10. The method of claim 9, wherein if the aerodynamic component is a fixed stabilizer, its local angle of attack is
Figure FDA0002651181400000041
Local sideslip angle
Figure FDA0002651181400000042
If the pneumatic part is a control surface, its deflection angle isiCorrecting the local angle of attack to
Figure FDA0002651181400000043
Local sideslip angle
Figure FDA0002651181400000044
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