CN110874055A - Prediction and control method for hypersonic aircraft separation process under action of two-phase flow field - Google Patents
Prediction and control method for hypersonic aircraft separation process under action of two-phase flow field Download PDFInfo
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Abstract
The method for predicting and controlling the separation process of the hypersonic aircraft under the action of the two-phase flow field is realized by the following steps: dispersing a flow field of the hypersonic aircraft by adopting an overlapping grid method; the fluid in the reverse rocket is equivalent to air, equivalent correction is carried out on the kinetic energy and momentum of the fluid, and an N-S equation of the equivalent flow field is established; performing CFD calculation of the flow field according to the established N-S equation of the equivalent flow field to obtain aerodynamic force and moment applied to the separation body in the separation process; solving the separation motion under different working conditions by using the obtained aerodynamic force and moment applied to the separation body to obtain the mass center motion and the attitude motion parameters of the separation body; the separation body comprises a precursor and a rear body.
Description
Technical Field
The invention belongs to the field of aircraft separation design, and relates to a high-precision simulation analysis method for a hypersonic aircraft separation process in an atmosphere.
Background
The hypersonic aircraft is gradually becoming a research hotspot in the fields of space transportation, military attack and defense game and the like by virtue of unique advantages. Because the hypersonic flight in the atmosphere, the separation process is influenced by factors such as extremely harsh environment, large dynamic pressure, jet flow interference and the like, so that the separation flow field is extremely complex and is coupled with the separation motion, and the separation process is difficult to predict. In the traditional separation design method, a steady aerodynamic data interpolation table is used as a basis for separation simulation calculation, and the requirement of reliable separation design of the type of flight cannot be met.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the invention provides a method for accurately predicting and actively controlling the separation process of a hypersonic statically unstable aircraft under the action of a two-phase flow field, and aims to solve the problem of separation design under the action of complex aerodynamic interference of the hypersonic aircraft in an atmosphere.
The technical solution of the invention is as follows: the method for predicting the separation process of the hypersonic aircraft under the action of the two-phase flow field is realized by the following steps:
dispersing a flow field of the hypersonic aircraft by adopting an overlapping grid method;
the fluid in the reverse rocket is equivalent to air, equivalent correction is carried out on the kinetic energy and momentum of the fluid, and an N-S equation of the equivalent flow field is established;
performing CFD calculation of the flow field according to the established N-S equation of the equivalent flow field to obtain aerodynamic force and moment applied to the separation body in the separation process;
solving the separation motion under different working conditions by using the obtained aerodynamic force and moment applied to the separation body to obtain the mass center motion and the attitude motion parameters of the separation body;
the separation body comprises a precursor and a rear body.
Further, the establishment of the N-S equation of the equivalent post-flow field is realized by the following steps:
simulating a flow field inside the spray pipe by adopting a multi-component N-S equation to obtain momentum and kinetic energy of a multi-component fluid at an outlet of the spray pipe;
secondly, obtaining air flow parameters at the outlet of the spray pipe by adopting the physical properties of air according to the fluid momentum kinetic energy expression and the momentum and kinetic energy of the multi-component fluid at the outlet of the spray pipe obtained in the first step;
thirdly, obtaining the simplified shape of the throat part of the nozzle and the flow parameters of the throat part according to the one-dimensional isentropic relation;
fourthly, modifying the throat expansion angle according to the Plante Meier expansion relational expression to ensure that the simplified gas expansion angle is consistent with the flow of multiple components;
and fifthly, establishing an N-S equation of the equivalent rear flow field by adopting the simplified throat shape, the modified nozzle expansion angle and the flow parameters of the throat.
Further, the separation motion under different working conditions is solved to obtain the parameters of the mass center motion and the attitude motion of the separation body, and the separation simulation prediction is carried out by adopting a flow field CFD and six-degree-of-freedom motion coupling calculation method.
Further, the coupling calculation discretizes the whole separation process into a plurality of time steps, and the specific steps are as follows:
(1) solving the flow field by using an N-S equation of the equivalent rear flow field on the current time step to obtain pneumatic force and moment acting on the front body and the rear body;
(2) substituting the pneumatic force and moment on the forebody and the rear body obtained at the current time step and other external forces into a six-degree-of-freedom rigid body motion equation together, and obtaining the displacement and posture change of the forebody and the rear body in the time step delta t to obtain the position and posture of the forebody and the rear body at the next time step;
(3) and (3) aiming at the new positions and postures of the precursor and the afterbody, obtaining new flow field dispersion by using the overlapped grids, turning to the step (1), and circulating the steps until the separation is finished.
Further, after the positions and postures of the front body and the rear body are obtained through calculation in the step (2), whether the front body and the rear body are safely separated or not is judged, and if the front body and the rear body collide or interfere, the current simulation is ended; otherwise, executing step (3).
Furthermore, the time step delta t is less than 5 ms.
A method for controlling the separation process of a hypersonic aircraft under the action of a two-phase flow field is characterized in that in the separation process, the pitch angle, the yaw angle and the roll angle are respectively and actively controlled by adopting a proportional feedback control law according to the centroid motion and the attitude motion parameters of a separation body calculated in real time.
Further, the control rudder deflection angle determined by adopting the proportional feedback control law is as follows:
wherein:
Dψ、Dγcontrolling rudder deflection angles for pitching, yawing and rolling respectively;
for the current pitch angle and angular velocity,
for the desired pitch angle and angular velocity,
a pitch angle feedback coefficient and an angular velocity feedback coefficient; psiu，
For the current yaw angle and angular velocity, #q，
For desired yaw angle and angular velocity, Kψ，CψThe yaw angle feedback coefficient and the angular speed feedback coefficient are obtained; gamma rayu，
For the current roll angle and angular velocity, gammaq，
For desired roll angle and angular velocity, Kγ，CγThe roll angle feedback coefficient and the angular velocity feedback coefficient.
Further, in the above-mentioned case,
Kψ、Kγthe value range of (1) is 2-4.
Further, in the above-mentioned case,
Cψ、Cγthe value range of (A) is 0.2-0.5.
Compared with the prior art, the invention has the beneficial effects that:
when the hypersonic aircraft is separated in the atmosphere, the hypersonic aircraft faces extremely harsh environment, large dynamic pressure, jet flow interference and other factors, so that the separation flow field is extremely complex, on one hand, the complex flow field causes the change rule of pneumatic force and moment acting on the aircraft to be complex, and the traditional pneumatic data interpolation table in a steady state cannot accurately obtain the real pneumatic force and moment acting on the aircraft; on the other hand, the aircraft is influenced by aerodynamic force and moment in the separation process, the attitude changes violently, the change of the attitude can cause the change of force and moment acting on the aircraft, the strong coupling relation exists between the attitude motion and the flow field of the aircraft in the separation process, and the separation motion process cannot be accurately predicted by the traditional calculation method. Therefore, the invention provides a two-phase flow field equivalent method, solves the problem of accurate prediction of pneumatic power and moment in the separation process, provides a flow field CFD and six-degree-of-freedom motion coupling calculation method based on overlapped grids, solves the problem of strong coupling between attitude motion and a flow field of an aircraft in the separation process, greatly improves the calculation precision compared with the traditional design method, and simultaneously provides a separation process active control method based on a proportional feedback control law on the basis of the result obtained by the method aiming at the characteristic of violent attitude change in the separation process of the aircraft, thereby solving the problem of reliable separation design of the aircraft.
The invention is successfully applied to 1: the scheme of the 1-plane symmetric boosting gliding missile verifies the flight test, the flight test is successful, and the accuracy of the method is effectively verified.
Drawings
FIG. 1 is a schematic diagram of an example of an overlapping grid of the present invention;
FIG. 2 is a flow chart of the N-S equation of the equivalent post-flow field of the present invention;
FIG. 3 is a schematic diagram comparing the whole separation solution propulsion process with the conventional calculation method;
FIG. 4 is a schematic diagram of the method of the present invention.
Detailed Description
The invention is described in detail below with reference to the figures and examples.
1. Mesh establishment
And establishing a partition overlapping grid required by the pneumatic CFD calculation according to the geometrical information of the aircraft and the separation body.
The partition overlapping mesh method is a discrete method of structure space. The method divides a complex flow area into sub-areas with simpler geometric boundaries, the computational grids in each sub-area are independently generated and have an overlapping or nested relation with each other, and flow field information is matched and coupled at the boundary of the overlapping area through interpolation. The overlapped grid method can conveniently process various complex flow field structures, is particularly suitable for flow field calculation of complex shape streaming and relative motion of objects, and after generation of each sub-grid is finished, the grid moves along with the objects without manual intervention.
The invention adopts an overlapping grid method as a flow field discrete method for CFD (computational fluid dynamics) calculation of the separation body aerodynamic characteristics of the hypersonic aircraft, and the overlapping grid is shown in figure 1.
2. Two-phase flow field equivalence
Aerodynamic force and moment applied to the aircraft in the separation process can be obtained by solving an N-S equation. The N-S equation has strong nonlinearity, no analytic solution exists at present, and a computer can be used for solving the numerical solution of the N-S equation under the initial condition and the boundary condition by a numerical calculation method. Namely, the finite difference is used for replacing the differential approximately, so that the N-S equation is converted into an algebraic equation, and the numerical solution is programmed.
And the flow field calculation adopts a high-precision numerical discrete format and a turbulence model, and the three-dimensional NS control equation set is solved. The system of equations in its conservative form is as follows:
where U is a conservative variable, E, F and G are convection terms in the x, y and z directions, respectively, Ev、FvAnd GvThe viscosity terms in the x, y and z directions, respectively.
The hypersonic speed aircraft usually generates required separation active impulse by means of a separation rocket in the separation process, the jet flow of the separation rocket is usually high-temperature two-phase flow, and the jet flow direction is generally opposite to the incoming flow direction of the aircraft, so that the flow field of the aircraft is very complicated. The accuracy of the flow field simulation of the thrust-back nozzle is the key for ensuring the whole separation simulation, and the simulation is necessary. In order to simplify the complexity of calculation and shield larger errors of pneumatic characteristic simulation caused by two-phase flow calculation deviation, the fluid in the reverse rocket is equivalent to air, equivalent correction is carried out on the kinetic energy and momentum of the fluid, and the CFD simulation calculation of a single component is realized, and the specific method comprises the following steps:
the throat of the jet pipe is endowed with engine throat parameters, the gamma of the outlet of the reverse thrust rocket is endowed through the whole flow field, the result obtained by calculating the gamma number is used as a standard value, the nozzle parameters (with dimension and area weighted average) of the jet pipe are obtained by calculation, and the nozzle parameters are compared with the throat parameters obtained by calculating the actual inlet conditions. The nozzle is modified by three aspects: and (3) modifying the expansion angle, modifying the throat and changing the entrance condition of the throat to obtain a reasonable jet flow simulation scheme. The calculation flow is shown in fig. 2, and may be specifically summarized as the following steps:
simulating a flow field inside the spray pipe by adopting a multi-component N-S equation to obtain momentum and kinetic energy of a multi-component fluid at an outlet of the spray pipe;
secondly, obtaining air flow parameters at the outlet of the spray pipe by adopting the physical properties of air according to the fluid momentum kinetic energy expression and the momentum and kinetic energy of the multi-component fluid at the outlet of the spray pipe obtained in the first step;
thirdly, obtaining the simplified shape of the throat part of the nozzle and the flow parameters of the throat part according to the one-dimensional isentropic relation;
fourthly, modifying the throat expansion angle according to the Plante Meier expansion relational expression to ensure that the simplified gas expansion angle is consistent with the flow of multiple components;
and fifthly, establishing an N-S equation of the equivalent rear flow field by adopting the simplified throat shape, the modified nozzle expansion angle and the flow parameters of the throat.
3. Flow field CFD and six-degree-of-freedom motion coupling calculation
a. Separate motion rigid body dynamics computation
(1) Movement of the center of mass of the separation body
The dynamic equation of the center of mass of the separation body under the separation coordinate system can be expressed as
Wherein Ω and V respectively represent a displacement motion velocity vector and a rotation angular velocity vector of the separation coordinate system with respect to the separator.
(2) Rotation of the separating body about the centre of mass
The equations describing the rotational motion of the separation body around the center of mass in the separation coordinate system can be expressed as:
wherein, IijI, j ═ x, y, z respectively represent the moment of inertia of the separation body about each axis of its elastic coordinate system; ω x, ω y, ω z represent the angular velocity component of the rotation of the separation body about its axis of the body coordinate system, respectively.
(3) Determination of the attitude angle of a separating body
Attitude angle of separated body
Gamma and the rotating angular speed of the separating body in the elastic body coordinate system have the following relation:
wherein the content of the first and second substances,
is the pitch angle of the wheels,
is the yaw angle and gamma is the roll angle.
According to the dynamic model, the separation motion under different working conditions can be solved, and the centroid motion and the attitude motion parameters of the separation body are obtained.
b. Separation simulation method for flow field CFD and six-degree-of-freedom motion coupling calculation
For hypersonic aircraft separation, a flow field is related to the motion state (separation distance, speed and the like) of the aircraft, the motion state of the aircraft is related to the aerodynamic force and moment of the process, and the flow field and the moment are closely coupled, so that the separation simulation analysis is carried out by adopting a flow field CFD and six-degree-of-freedom motion coupling calculation method, and the prediction precision is improved.
The whole separation solving propulsion process is compared with the traditional calculation method, and the method is shown in figure 3.
In the solution of the invention, the whole separation process is divided into a plurality of time periods, and the flow field is finely solved at each time step to obtain the pneumatic force and the moment acting on the front body and the rear body; further, substituting the displacement and the posture change of the front body and the rear body in a time step delta t (generally less than 5ms) together with other external forces into a six-degree-of-freedom rigid body motion equation to obtain the position and the posture of the next moment; then generating a new grid for the shape of the precursor and the afterbody at the new position and the new posture to carry out flow field calculation; and circulating the steps until the separation is finished. The method considers the coupling relation of jet flow, a separation flow field and two-body motion, and can accurately predict the control starting posture boundary.
In order to improve the reliability of the method, whether the front body and the rear body are safely separated or not is judged after the positions and the postures of the front body and the rear body are obtained through calculation, if the front body and the rear body collide or interfere with each other, the current simulation is ended, and if not, the execution is continued.
4. Separate disturbance active control
The hypersonic aircraft is separated in the atmosphere, usually due to the influence of high dynamic pressure and jet flow interference, the pneumatic interference has large influence on the attitude of the aircraft and is unfavorable for starting and controlling the separated aircraft, and therefore the invention further provides an active control method for the separation interference.
And respectively adopting proportional feedback control laws for the pitch angle, the yaw angle and the roll angle, namely controlling the rudder deflection angle as follows:
wherein:
Dψ、Dγcontrolling rudder deflection angles for pitching, yawing and rolling respectively;
for the current pitch angle and angular velocity,
for the desired pitch angle and angular velocity,
a pitch angle feedback coefficient and an angular velocity feedback coefficient; psiu，
For the current yaw angle and angular velocity, #q，
For desired yaw angle and angular velocity, Kψ，CψThe yaw angle feedback coefficient and the angular speed feedback coefficient are obtained; gamma rayu，
For the current roll angle and angular velocity, gammaq，
For desired roll angle and angular velocity, Kγ，CγThe roll angle feedback coefficient and the angular velocity feedback coefficient.
Kψ、KγThe value range of (A) is 2-4,
Cψ、Cγthe value range of (A) is 0.2-0.5.
And (4) performing feedback control of the starting control rudder deflection according to the attitude interference calculation result in the separation process, and adjusting separation design parameters (such as time sequence, criterion and the like) to perform the process again until the design requirements are met if the design requirements cannot be met.
The above process can be implemented in engineering by using the framework shown in fig. 4, and the aircraft profile module and the jet flow equivalent module are original inputs; the grid file standard control interface and the solver interface are interface modules and are used for processing the input information into standard interface format files which can be identified by the solver; the CFD solver and the six-degree-of-freedom calculation module are solver modules and are used for carrying out CFD solver and six-degree-of-freedom calculation; the separating body position and posture result module is an output and post-processing module, outputs a calculation result and performs interference, collision inspection and other work; and the other modules are preprocessing modules used for converting the originally input physical conditions into corresponding calculation input conditions.
The invention has not been described in detail in part of the common general knowledge of those skilled in the art.
Claims (10)
1. The method for predicting the separation process of the hypersonic aircraft under the action of the two-phase flow field is characterized by being realized in the following mode:
dispersing a flow field of the hypersonic aircraft by adopting an overlapping grid method;
the fluid in the reverse rocket is equivalent to air, equivalent correction is carried out on the kinetic energy and momentum of the fluid, and an N-S equation of the equivalent flow field is established;
performing CFD calculation of the flow field according to the established N-S equation of the equivalent flow field to obtain aerodynamic force and moment applied to the separation body in the separation process;
solving the separation motion under different working conditions by using the obtained aerodynamic force and moment applied to the separation body to obtain the mass center motion and the attitude motion parameters of the separation body;
the separation body comprises a precursor and a rear body.
2. The method of claim 1, wherein: the establishment of the N-S equation of the equivalent flow field is realized by the following method:
simulating a flow field inside the spray pipe by adopting a multi-component N-S equation to obtain momentum and kinetic energy of a multi-component fluid at an outlet of the spray pipe;
secondly, obtaining air flow parameters at the outlet of the spray pipe by adopting the physical properties of air according to the fluid momentum kinetic energy expression and the momentum and kinetic energy of the multi-component fluid at the outlet of the spray pipe obtained in the first step;
thirdly, obtaining the simplified shape of the throat part of the nozzle and the flow parameters of the throat part according to the one-dimensional isentropic relation;
fourthly, modifying the throat expansion angle according to the Plante Meier expansion relational expression to ensure that the simplified gas expansion angle is consistent with the flow of multiple components;
and fifthly, establishing an N-S equation of the equivalent rear flow field by adopting the simplified throat shape, the modified nozzle expansion angle and the flow parameters of the throat.
3. The method of claim 1, wherein: and solving the separation motion under different working conditions to obtain the parameters of the mass center motion and the attitude motion of the separation body, and performing separation simulation prediction by adopting a flow field CFD and six-degree-of-freedom motion coupling calculation method.
4. The method of claim 3, wherein: the coupling calculation discretizes the whole separation process into a plurality of time steps, and the specific steps are as follows:
(1) solving the flow field by using an N-S equation of the equivalent rear flow field on the current time step to obtain pneumatic force and moment acting on the front body and the rear body;
(2) substituting the pneumatic force and moment on the forebody and the rear body obtained at the current time step and other external forces into a six-degree-of-freedom rigid body motion equation together, and obtaining the displacement and posture change of the forebody and the rear body in the time step delta t to obtain the position and posture of the forebody and the rear body at the next time step;
(3) and (3) aiming at the new positions and postures of the precursor and the afterbody, obtaining new flow field dispersion by using the overlapped grids, turning to the step (1), and circulating the steps until the separation is finished.
5. The method of claim 4, wherein: after the positions and postures of the front body and the rear body are obtained through calculation in the step (2), whether the front body and the rear body are safely separated or not is judged, and if the front body and the rear body collide or interfere, the current simulation is ended; otherwise, executing step (3).
6. The method of claim 3, wherein: the time step delta t is less than 5 ms.
7. The method for controlling the separation process of the hypersonic aircraft under the action of the two-phase flow field is characterized by comprising the following steps: and in the separation process, the pitch angle, the yaw angle and the roll angle are actively controlled by adopting a proportional feedback control law according to the real-time calculated mass center motion and attitude motion parameters of the separation body.
8. The method of claim 7, wherein: the control rudder deflection angle determined by adopting a proportional feedback control law is as follows:
wherein:
Dψ、Dγcontrolling rudder deflection angles for pitching, yawing and rolling respectively;
for the current pitch angle and angular velocity,
for the desired pitch angle and angular velocity,
a pitch angle feedback coefficient and an angular velocity feedback coefficient; psiu，
For the current yaw angle and angular velocity, #q，
For desired yaw angle and angular velocity, Kψ，CψThe yaw angle feedback coefficient and the angular speed feedback coefficient are obtained; gamma rayu，
For the current roll angle and angular velocity, gammaq，
For desired roll angle and angular velocity, Kγ，CγThe roll angle feedback coefficient and the angular velocity feedback coefficient.
9. The method of claim 8, wherein:
Kψ、Kγthe value range of (1) is 2-4.
10. The method of claim 8, wherein:
Cψ、Cγthe value range of (A) is 0.2-0.5.
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| CN112068581B (en) * | 2020-09-11 | 2023-11-17 | 中国运载火箭技术研究院 | Aircraft composite control method, control device and storage medium |
| CN112182771A (en) * | 2020-10-11 | 2021-01-05 | 中国运载火箭技术研究院 | Data processing method based on numerical simulation, storage medium and electronic device |
| CN112182771B (en) * | 2020-10-11 | 2022-08-05 | 中国运载火箭技术研究院 | Data processing method based on numerical simulation, storage medium and electronic device |
| CN112528420A (en) * | 2020-12-25 | 2021-03-19 | 中国空气动力研究与发展中心计算空气动力研究所 | Dynamic boundary condition switching method for jet flow time sequence control simulation |
| CN113886942A (en) * | 2021-09-01 | 2022-01-04 | 北京机电工程研究所 | Numerical simulation method for aircraft hood hinge constraint ejection separation |
| CN113867381A (en) * | 2021-12-02 | 2021-12-31 | 中国空气动力研究与发展中心计算空气动力研究所 | Aircraft attitude control method |
| CN113867381B (en) * | 2021-12-02 | 2022-02-22 | 中国空气动力研究与发展中心计算空气动力研究所 | Aircraft attitude control method |
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