Rapid prediction method for supersonic rotation dynamic derivative of aircraft

Abstract

The invention provides a rapid prediction method for a supersonic rotation dynamic derivative of an aircraft. The method comprises the following steps: firstly, solving a steady state aerodynamic torque coefficient Cm0 of the aircraft under a given working condition; secondly, obtaining a corresponding aerodynamic moment coefficient Cm of the aircraft at the additional rotating speed q under the same working condition through quasi-steady calculation, and further obtaining the disturbance quantity of the aerodynamic moment; and finally, directly obtaining the rotation dynamic derivative based on the definition # imgabs0 # of the rotation dynamic derivative. According to the method, the rotating dynamic derivative of the supersonic aircraft can be obtained on the basis that only quasi-steady solution is carried out, multiple times of unsteady forced vibration calculation in the design process are avoided, meanwhile, high prediction precision is achieved, and technical support can be provided for rapid evaluation of the ultrasonic dynamic derivative of the aircraft.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft aerodynamic performance prediction, and in particular to a method for quickly predicting supersonic rotation dynamic derivatives of an aircraft. Background Art

[0002] Dynamic derivatives are not only important parameters for analyzing the flight stability and flight quality of aircraft, but also an important basis for determining the control gain in the design of aircraft attitude and trajectory control systems. The rotational dynamic derivatives required in the analysis of aircraft stability are generally obtained by designing a forced vibration method for displacement motion.

[0003] In order to obtain a stable periodic unsteady flow field, a large amount of computing resources and computing time are required, which cannot meet the needs of quickly evaluating the dynamic characteristics of the aircraft. The existing engineering methods have problems such as low prediction accuracy and lack of verification. Summary of the invention

[0004] The purpose of the present invention is to provide a method for quickly predicting the supersonic rotational dynamic derivatives of an aircraft, aiming to reduce the simulation difficulty and calculation cycle, and provide technical support for the rapid evaluation of the dynamic stability characteristics of the aircraft.

[0005] According to the purpose of the present invention, the present invention provides a method for quickly predicting the supersonic rotational dynamic derivative of an aircraft, comprising the following steps:

[0006] (1) Solve the steady-state aerodynamic moment coefficient C of the aircraft under given working conditions m0 ;

[0007] (2) The corresponding aerodynamic moment coefficient C is obtained by quasi-steady calculation when the aircraft has the same additional rotation speed q under the same working condition m , and then obtain the aerodynamic disturbance;

[0008] (3) Using the aerodynamic moment coefficient C determined in steps (1) and (2) m0 and C m , based on the definition of derivative Directly obtain the rotational derivative.

[0009] Furthermore, in step (1), C m0 It is the rolling moment coefficient Cmx, yaw moment coefficient Cmy or pitch moment coefficient Cmz around the center of mass of the aircraft.

[0010] Furthermore, in step (1), the steady-state aerodynamic moment coefficient C m0 The NS equation is used for solution.

[0011] Furthermore, in step (1), the given operating conditions are a Mach number range of 1.5 to 40, a flight angle of attack range of 0° to 360°, and a flight sideslip angle range of 0° to 360°.

[0012] Furthermore, in step (2), the disturbance motion speed q is the rolling angular velocity ω around the center of mass of the aircraft x , yaw angular velocity ω y Or the pitch angular velocity ω z .

[0013] Furthermore, in step (2), C m Add the rolling angular velocity ω to the aircraft x Rolling moment coefficient Cmx around the center of mass, additional yaw angular velocity ω y Yaw moment coefficient Cmy around the center of mass or additional pitch angular velocity ω z The pitching moment coefficient Cmz about the center of mass at .

[0014] Furthermore, in step (3), the rotational derivative is the rolling derivative Cmx around the center of mass of the aircraft ωx , yaw dynamic derivative Cmy ωy Or pitch derivative Cmy ωz .

[0015] Furthermore, in step (2), the quasi-steady-state calculation adopts a solution perturbation method based on the Euler equation, and through R invariant perturbation, a perturbation relationship is constructed for the normal velocity of the object surface caused by the rotational disturbance, and the rotational derivative is quickly calculated.

[0016] Furthermore, in step (2), the quasi-steady calculation adopts a fast calculation method of unsteady aerodynamic force based on local flow piston theory, combined with the downwash velocity of surface deformation or vibration, to calculate the additional disturbance term of the aerodynamic force.

[0017] Furthermore, in step (3), the calculation of the rotational dynamic derivative takes into account the dynamic stability characteristics of the aircraft under supersonic and hypersonic conditions, and is suitable for rapid prediction of dynamic derivatives of aircraft with complex aerodynamic shapes.

[0018] The technical solution of the present invention, the rapid aircraft prediction method, can quickly obtain the aircraft rotational dynamic derivatives based on only calculating the steady and quasi-steady aerodynamic moments, avoiding multiple unsteady forced vibration calculations during the design process, and at the same time has high prediction accuracy, which can provide technical support for the rapid prediction of supersonic aircraft rotational dynamic derivatives. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0020] Figure 1 is a flow chart of an embodiment of the present invention;

[0021] Figure 2 A schematic diagram of a calculation model according to an embodiment of the present invention; DETAILED DESCRIPTION

[0022] The technical solution of the present invention will be clearly and completely described below in conjunction with the embodiments. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0023] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "clockwise", "counterclockwise" and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as limiting the present invention.

[0024] In addition, the terms "first" and "second" are only used for descriptive purposes and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Therefore, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, "multiple" means two or more, unless otherwise clearly and specifically defined. In addition, the terms "installed", "connected" and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal connection of two components. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0025] Example 1

[0026] like Figure 1-Figure 2 As shown,

[0027] A method for quickly predicting supersonic rotational dynamic derivatives of an aircraft comprises the following steps:

[0028] Step 1: Calculate the aerodynamic moment coefficient C of the aircraft in steady state m0 ;

[0029] Step 2: Obtain the corresponding aerodynamic torque coefficient C when the aircraft has the same additional rotation speed q under the same working condition through quasi-steady calculation m , and then obtain the aerodynamic disturbance;

[0030] Step 3, based on the definition of derivatives Directly obtain the rotational derivative.

[0031] More specifically:

[0032] (1) In step 1, solve the steady-state aerodynamic moment coefficient C of the aircraft under given working conditions. m0 ;

[0033] Specifically, the flight Mach number range is 1.5 to 40, the flight angle of attack range is 0° to 360°, and the flight sideslip angle range is 0° to 360°.

[0034] C m0 The value can be the rolling moment coefficient Cmx, the yaw moment coefficient Cmy or the pitch moment coefficient Cmz around the center of mass of the aircraft, which is obtained by solving the steady-state NS equation.

[0035] (2) In step 2, the corresponding aerodynamic moment coefficient C is obtained by quasi-steady calculation when the aircraft has the same working condition and additional rotation speed q m , and then obtain the aerodynamic disturbance;

[0036] Specifically, the disturbance motion speed q can be the rolling angular velocity ω around the center of mass of the aircraft x , yaw angular velocity ω y Or the pitch angular velocity ω z ;

[0037] C m Add the rolling angular velocity ω to the aircraft x Rolling moment coefficient Cmx around the center of mass, additional yaw angular velocity ω y Yaw moment coefficient Cmy around the center of mass or additional pitch angular velocity ω z The pitching moment coefficient Cmz about the center of mass at .

[0038] The quasi-steady calculation adopts the solution perturbation method based on the Euler equation. Through the R invariant perturbation, the perturbation relationship of the normal velocity of the object surface caused by the rotational disturbance is constructed, and the rotational derivative is quickly calculated.

[0039] Quasi-steady calculations can also use a fast calculation method for unsteady aerodynamics based on local flow piston theory, combined with the downwash velocity of surface deformation or vibration, to calculate additional disturbance terms of the aerodynamics.

[0040] (3) In step 3, the aerodynamic moment coefficient C determined in steps 1 and 2 is used. m0 and C m , based on the definition of derivative Directly obtain the rotational derivative.

[0041] Specifically, the rotational derivative C mq is the rolling derivative Cmx around the center of mass of the vehicle ωx , yaw dynamic derivative Cmy ωy Or pitch derivative Cmy ωz The calculation of the rotational dynamic derivative takes into account the dynamic stability characteristics of the aircraft under supersonic and hypersonic conditions, and is suitable for the rapid prediction of the dynamic derivative of aircraft with complex aerodynamic shapes.

[0042] Example 2

[0043] like Figure 2 As shown, the specific solution example of the rapid prediction of the supersonic dynamic derivative of the aircraft in this embodiment is as follows:

[0044] This example uses an X-shaped rudder missile shape. Figure 2 The main dimensions of the calculation model for this example are given, with a reference length of 2.4m and a reference area of ​​0.025m. 2 , the aircraft center of mass coefficient is (0.52, 0, 0).

[0045] The Mach numbers for comparative calculation and analysis are 3, 4, and 5, respectively, corresponding to a flight altitude range of 0 to 20 km. Given a roll angular velocity of 20 rad / s, a comparative analysis is performed on the roll derivatives for the states of 0° angle of attack and 0° sideslip angle.

[0046] Table 1 lists the steady-state rolling moment coefficient, the quasi-steady rolling moment coefficient at a rolling angular velocity of 20 rad / s, and the rolling rotational derivative values ​​determined by the example method of the present invention. Table 2 lists the rolling rotational derivative values ​​determined by the free oscillation unsteady numerical simulation method.

[0047] It can be found that the method for rapid prediction of supersonic rotational derivatives of aircraft according to the embodiment of the present invention can obtain rotational derivative results close to those of numerical methods. The computational efficiency of this method is more than one order of magnitude higher than that of the free vibration numerical simulation method, and is very suitable for rapid calculation of large-scale aerodynamic data required in multiple rounds of iterations during the aircraft design phase.

[0048] Table 1:

[0049] Ma Cmx0 <![CDATA[Dimensionless ω x > Cmx <![CDATA[Cmx / ω x ]]> 3 1.0341E-04 2.0310E-02 3.7994E-04 -1.3615E-02 4 7.3585E-05 1.5233E-02 2.3089E-04 -1.0327E-02 5 5.2260E-05 1.2186E-02 1.4890E-04 -7.9301E-03

[0050] Table 2:

[0051] Ma <![CDATA[Cmx / ω x ]]> 3 -1.557E-02 4 -1.230E-02 5 -9.152E-03

[0052] The rapid prediction method of the aircraft of the present invention can quickly obtain the rotational dynamic derivatives of the aircraft based on the calculation of only steady and quasi-steady aerodynamic moments, avoiding multiple unsteady forced vibration calculations during the design process. At the same time, it has high prediction accuracy and can provide technical support for the rapid prediction of the rotational dynamic derivatives of supersonic aircraft.

[0053] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or replace some or all of the technical features therein with equivalents. However, these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for rapidly predicting supersonic rotational derivatives of an aircraft, characterized in that: The following steps are involved: S1. Solve the aerodynamic moment coefficient C of the aircraft under given working conditions m0 ; S2. Obtain the corresponding aerodynamic moment coefficient C when the aircraft has the same additional rotation speed q under the same working condition through quasi-steady calculation. m , and then obtain the aerodynamic disturbance; S3, aerodynamic moment coefficient C determined by S1 and S2 m0 and C m , based on the definition of derivative Directly obtain the rotational derivative.

2. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: S1, C m0 It is the rolling moment coefficient Cmx, yaw moment coefficient Cmy or pitch moment coefficient Cmz around the center of mass of the aircraft.

3. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S1, the steady-state aerodynamic moment coefficient C m0 The NS equation is used for solution.

4. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S1, the given operating conditions are Mach number range 1.5 ~ 40, flight angle of attack range 0 ° ~ 360 °, and flight sideslip angle range 0 ° ~ 360 °.

5. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S2, the disturbance motion speed q is the rolling angular velocity ω around the center of mass of the aircraft x , yaw angular velocity ω y Or the pitch angular velocity ω z .

6. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: S2, C m Add the rolling angular velocity ω to the aircraft x Rolling moment coefficient Cmx around the center of mass, additional yaw angular velocity ω y Yaw moment coefficient Cmy around the center of mass or additional pitch angular velocity ω z The pitching moment coefficient Cmz about the center of mass at .

7. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S3, the rotational derivative is the rolling derivative Cmx around the center of mass of the aircraft. ωx , yaw dynamic derivative Cmy ωy Or pitch derivative Cmy ωz .

8. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S2, the quasi-steady-state calculation adopts a solution perturbation method based on the Euler equation, and through R invariant perturbation, a perturbation relationship is constructed for the normal velocity of the object surface caused by the rotational disturbance, and the rotational derivative is quickly calculated.

9. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S2, the quasi-steady calculation adopts a fast calculation method of unsteady aerodynamic force based on local flow piston theory, combined with the downwash velocity of surface deformation or vibration, to calculate the additional disturbance term of the aerodynamic force.

10. The method for rapid prediction of supersonic rotational derivatives of an aircraft according to claim 1, characterized in that: In S3, the calculation of the rotational dynamic derivative takes into account the dynamic stability characteristics of the aircraft under supersonic and hypersonic conditions, and is suitable for rapid prediction of dynamic derivatives of aircraft with complex aerodynamic shapes.

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