CN111332491B - Method for improving pneumatic servo elastic stability - Google Patents
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Abstract
The invention discloses a design method for improving pneumatic servo elastic stability, which adopts a full-aircraft ground resonance test or a full-aircraft ground structure modal coupling test data design optimization technology to correct an elastic model and a transfer function of an aircraft body, and then adopts a flight control law dynamic parameter adjusting technology to improve the pneumatic servo elastic stability; the invention can well give consideration to the coupling relation among aerodynamic force, structural elasticity and a control system, thereby realizing the purpose of meeting the flight quality requirement, the stability margin requirement of the control system and the stability requirement of pneumatic servo elasticity in a full envelope range, and being applied and verified in the unmanned aerial vehicle with multiple control surfaces. The problem of a kind of many control surfaces unmanned aerial vehicle pneumatic servo elastic stability be difficult to control is solved.
Description
Technical Field
The invention belongs to the field of aviation flight control and pneumatic servo elasticity combined design, and particularly relates to a method for improving the stability of pneumatic servo elasticity.
Background
The pneumatic servo elasticity refers to the characteristic of mutual coupling of aircraft structure dynamics, unsteady aerodynamic force and a servo control system (including an inertial set, a stability augmentation control law and a steering engine). The sensor on the airframe receives not only the rigid motion signal of the aircraft, but also the elastic vibration signal of the structure, the signal is converted into the deflection signal of the control surface through the servo control system, thus forming the pneumatic control force; the control force, in turn, affects the structure elastic vibration. The structure and the control are mutually fed back to form a pneumatic servo elastic closed loop system.
The presence of such an unfavorable coupling between the structure, the pneumatic and the control system can lead to a deterioration in the performance of the closed-loop system and even to a loss of stability. And as the aircraft develops towards the goals of light structure, higher speed and better performance, the problem is increasingly prominent, and the flight safety and performance are directly influenced. To ensure stability of an aircraft with a servo flight control system, the aero-servo elastic stability margin must meet aircraft strength and stiffness specifications. Through more than sixty years of research, the pneumatic servo elasticity analysis, synthesis and test have made significant progress domestically and abroad. Over the last decade several new problems have emerged due to the development of non-conventional aircraft designs, imposing significant challenges on pneumatic servo research.
At present, the method for improving the pneumatic servo elasticity at home and abroad mainly adopts the methods of optimizing the position of a flight control sensor, optimizing a filter and the like. The layout design of the flight control sensors such as a rate gyro, an acceleration sensor, an attack angle sensor and the like is generally carried out in the early stage of the aircraft according to theoretical basis and past model experience. Pneumatic servo elastic motion and rigid motion coupling are eliminated, and the pneumatic servo elastic motion can be well restrained by adopting a wave trap design. However, the complex coupling among aerodynamic force, structural elasticity and a control system causes the requirement of improving the pneumatic servo elasticity to consider flight quality and requirements, stability margin requirement of the control system and stability requirement of the pneumatic servo elasticity at the same time, and is a complex multi-disciplinary optimization technical problem, so that the capability of improving the pneumatic servo elasticity by the traditional method is limited.
Disclosure of Invention
The invention provides a method for improving the pneumatic servo elastic stability, aiming at the problem of limited capability of improving the pneumatic servo elastic stability in the prior art.
The invention specifically comprises the following contents:
the invention relates to a method for improving pneumatic servo stability, which comprises the steps of firstly establishing an airplane body model of an airplane and an airplane link transfer function of the airplane body model, optimizing the airplane body model, and then improving pneumatic servo elastic stability by using a flight control law dynamic parameter adjusting technology.
In order to better implement the invention, further, the specific operations of improving the pneumatic servo elastic stability by using the flight control dynamic parameter adjusting technology are as follows: in the combined design of aeroelasticity and a flight control system, a rigid body mode and elastic mode separation technology is applied to carry out aeroelasticity servo stability analysis on an airplane system, calculate the frequency range of undamped oscillation of a long-period mode and a short-period mode, and simultaneously calculate the 1 st order elastic frequency of the airplane body structure of the airplane; the low elastic modal frequency is far higher than the long and short period modal frequency, namely the rigid body modal frequency is separated from the elastic modal frequency, and the problem of rigid-elastic coupling aeroelastic stability is not easy to occur.
In order to better realize the method, the machine body model is further linearized at each designed dynamic pressure point, cut-off frequency, phase angle margin, crossing frequency and amplitude margin are calculated by loop simulation, the robustness of the machine body model is analyzed, and whether the dynamic parameters of the control law in the full envelope range have consistent control quality is verified, so that the control law robustness and flight safety of the airplane are improved.
In order to better realize the method, further, whether amplitude margin of a longitudinal channel, a transverse channel and a course channel of the airplane in a full envelope range meets a condition of more than 6db or not and whether phase margin meets a condition of more than 45 degrees or not are checked through pneumatic servo elastic stability analysis; if not, the corresponding parameter adjustment is carried out through the flight control rhythm dynamic parameter adjustment technology.
In order to better realize the method, further, the optimization of the model needs to fully consider an aircraft structure finite element model, so that a full-aircraft ground resonance test under multipoint sinusoidal excitation is adopted to obtain data of the model in three aspects of generalized mass, generalized rigidity and structural damping coefficient, the data is used for correcting the full-aircraft finite element model, the modal, the slope and the control surface inertia coupling parameters of the position of the sensor are corrected through an open loop frequency response characteristic test, and then, whether the open loop amplitude-frequency curves of the control laws of a longitudinal channel, a transverse channel and a course channel after the aircraft correction are close to a test value or not is compared.
In order to better realize the method, data of the model in three aspects of generalized mass, generalized rigidity and structural damping coefficient are obtained by adopting a full-aircraft ground structure modal coupling test and are used for correcting a full-aircraft finite element model, the modal, the slope and the control surface inertial coupling parameters of the position of a sensor are corrected through an open-loop frequency response characteristic test to optimize and correct the aircraft finite element model, and then, whether the open-loop amplitude-frequency curves of the control law of a longitudinal channel, a transverse channel and a course channel after the aircraft is corrected are compared to be close to a test value.
In order to better implement the method, the natural frequency characteristics of the rigid body of the airframe structure of the airplane and the elastic frequency characteristics of the airframe are further analyzed, and the coupling of the rigid body mode and the elastic mode is avoided.
In order to better implement the method, further, an airplane link transfer function of the airframe model is a ratio of a signal received by the sensor to a deflection degree of a control plane.
Compared with the prior art, the invention has the following advantages and beneficial effects:
(1) the method of multidisciplinary combined pneumatic servo elastic stability can well give consideration to the coupling relation among aerodynamic force, structural elasticity and a control system;
(2) the machine body elastic model and the transfer function are corrected by using the data of the whole-machine ground resonance test and the whole-machine structure modal coupling test, so that the pneumatic servo elastic stability analysis is more accurate;
(3) by separating the rigid body mode from the elastic mode, the problem of coupling between the rigid body mode and the elastic mode of the airplane body in the design of the airplane is solved.
Drawings
FIG. 1 is a schematic diagram of a model correction principle;
FIG. 2 is a flow chart of dynamic parameter adjustment for pneumatic servo elastic stability;
FIG. 3 is a comparison of the open loop amplitude-frequency curves of the control law for the longitudinal channel;
FIG. 4 is a comparison of the open loop amplitude-frequency curves of the control law of the transverse channel;
FIG. 5 is a comparison of open loop amplitude-frequency curves of the course channel control law;
fig. 6 shows the variation of the magnitude margin of the course channel with the course control parameter krr of the drone.
Detailed Description
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it should be understood that the described embodiments are only a part of the embodiments of the present invention, and not all embodiments, and therefore should not be considered as a limitation to the scope of protection. All other embodiments, which can be obtained by a person skilled in the art without any inventive step based on the embodiments of the present invention, are within the scope of the present invention.
Example 1:
the invention relates to a method for improving pneumatic servo elastic stability, which is shown by combining with a figure 1, a figure 2, a figure 3, a figure 4, a figure 5 and a figure 6, taking a certain type of unmanned aerial vehicle as an example, and taking the principle as the figure 1, firstly establishing an organism elastic model and a transfer function, and carrying out model correction by combining with a whole-aircraft ground resonance test and a whole-aircraft structure modal coupling test; as shown in the flow chart of fig. 2, the influence of the parameters of the control law of the airplane on the stability and flight quality of the aerodynamic servo elasticity is researched, and the control law is dynamically adjusted and optimized on the basis, so that the airplane can simultaneously meet the requirements on the flight quality and the stability of the aerodynamic servo elasticity;
correcting the finite element model of the whole machine by using the ground resonance test of the whole machine; simulating the free horizontal flight state of the airplane, adopting a multipoint sine excitation phase resonance method, using generalized mass, generalized rigidity and structural damping coefficient obtained by testing to correct the model, performing difference analysis on the model calculation result and the test result, and reasonably correcting the parameters of the finite element model; the calculated values and the test values before and after the correction of the elastic vibration modal frequency of the sample unmanned aerial vehicle are shown in the following tables 1-4, and it can be seen that the model error is reduced after the correction;
TABLE 1 modified front and rear elastic vibration modal frequency contrast
TABLE 2 modified front and rear elastic vibration modal frequency contrast
TABLE 3 modified front and rear elastic vibration modal frequency contrast
TABLE 4 modified front and rear elastic vibration modal frequency contrast
Measuring the open-loop frequency response characteristic of the flight control system by using a full-aircraft ground structure modal coupling test, and correcting the modal and the slope of the position of the sensor and the inertial coupling parameter of the control surface; the sample unmanned aerial vehicle corrected longitudinal, transverse and course channel control law open-loop amplitude-frequency curves are shown in FIGS. 3-5;
analyzing the inherent frequency characteristic of the rigid body and the elastic frequency characteristic of the body of the unmanned aerial vehicle structure, and avoiding the coupling of the rigid body mode and the elastic mode as much as possible; calculating to obtain a sample unmanned aerial vehicle short-period modal undamped oscillation frequency of 0.22-0.7 Hz, long-period modal undamped oscillation frequency of 0.01-0.04 Hz, and organism structure 1-order elastic frequency of about 3.4 Hz; the low-elasticity modal frequency is about 5-15 times of the short-period modal frequency and is higher than the long-period modal frequency by two orders of magnitude, the rigid body modal frequency and the elasticity modal frequency are separated relatively, and the problem of rigid-elastic coupling aeroelasticity stability is not easy to occur;
the full-aircraft pneumatic servo elastic stability analysis of the sample unmanned aerial vehicle shows that the longitudinal and transverse channels meet the stability requirements that the amplitude margin is larger than 6db and the phase margin is larger than 45 degrees; but the amplitude margin of the course channel meets the requirement under low-speed pressure and does not meet the requirement under high-speed pressure;
analyzing the influence of unsatisfied loop control parameters on the pneumatic servo elastic stability, and carrying out sensitivity analysis on the sample unmanned aerial vehicle heading control parameter krr to obtain a heading channel amplitude margin follow-up dynamic pressure change curve of different unmanned aerial vehicle heading control parameters krr conditions in a design envelope, as shown in FIG. 6; as can be seen from the figure, the magnitude margin decreases as the value of the drone heading control parameter krr increases; in order to ensure that the amplitude margin is larger than 6db, the unmanned aerial vehicle heading control parameter krr can be subjected to dynamic parameter adjustment through pressure change, and the value of the unmanned aerial vehicle heading control parameter krr is not larger than 0.5 at a large dynamic pressure position;
and then analyzing the influence of the unsatisfied loop control parameters on the flight quality, and calculating the Holland rolling mode damping ratio corresponding to different unmanned aerial vehicle heading control parameters krr of the sample unmanned aerial vehicle as shown in the following tables 5-6:
TABLE 5 Holland roll mode damping ratio before dynamic parameter adjustment
TABLE 6 Holland roll mode damping ratio before dynamic parameter adjustment
It can be seen that the smaller the unmanned aerial vehicle heading control parameter krr is, the smaller the damping ratio is, when the unmanned aerial vehicle heading control parameter krr takes values of 0.8 and 1.0, the full envelope damping ratio is near the optimal damping ratio of 0.707, when the unmanned aerial vehicle heading control parameter krr takes values of 0.4 and 0.6, the full envelope damping ratio is both smaller than the optimal damping ratio of 0.707, and the rigid damping characteristic requirement of the unmanned aerial vehicle structure is not met, so that the heading control parameter krr of the full envelope unmanned aerial vehicle cannot be uniformly reduced to a certain smaller value; it can also be seen that the damping ratio is relatively ideal in the case of a small angle of attack, i.e., a large dynamic pressure, but the damping ratio is somewhat insufficient in the case of a large angle of attack, i.e., a small dynamic pressure, and therefore the dynamic parameter-adjusting scheme is determined as follows: as shown in table 7, the parameter of the unmanned aerial vehicle heading control parameter krr is adjusted according to the dynamic pressure, that is, the unmanned aerial vehicle heading control parameter krr is adjusted to be half at the large dynamic pressure point, and the unmanned aerial vehicle heading control parameters krr at the rest dynamic pressure points are only slightly adjusted to be small;
| parameter value | 2674 | 4961 | 8586 |
| Krr | 0.7 | 0.7 | 0.5 |
TABLE 7 dynamic pressure parameter modulation scheme
The damping ratio of the dutch roll mode after dynamic pressure parameter adjustment is shown in table 8, and it can be seen that the damping ratio of the dutch roll mode is reduced to some extent after the dynamic pressure parameter adjustment, but the integral damping ratio accords with the rigid damping characteristic of the unmanned aerial vehicle structure;
TABLE 8 Holland Rolling Mount Mode data after dynamic parameter adjustment
The results of the analysis of the aerodynamic servo elasticity stability of the airplane after the dynamic parameter adjustment are shown in the following tables 9 to 10, and it can be seen from the results that the aerodynamic servo elasticity stability margin in the full envelope range meets the requirements after the parameter adjustment.
TABLE 9 pneumatic servo elastic stability calculation State and results after dynamic parameter adjustment
TABLE 10 pneumatic servo elastic stability calculation State and results after dynamic parameter adjustment
The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention in any way, and all simple modifications and equivalent variations of the above embodiments according to the technical spirit of the present invention are included in the scope of the present invention.
Claims (7)
1. A method for improving the pneumatic servo elastic stability is characterized in that an airplane body model of an airplane and an airplane link transfer function of the airplane body model are established, the airplane body model is optimized, and then the pneumatic servo elastic stability is improved by using a flight control law dynamic parameter adjusting technology;
the specific operation of improving the pneumatic servo elastic stability by using the flight control dynamic parameter adjusting technology is as follows: in the combined design of aeroelasticity and a flight control system, a rigid body mode and elastic mode separation technology is applied to carry out aeroelasticity servo stability analysis on an airplane system, calculate the frequency range of undamped oscillation of a long-period mode and a short-period mode, and simultaneously calculate the 1 st order elastic frequency of the airplane body structure of the airplane; the low elastic modal frequency is far higher than the long and short period modal frequency, namely the rigid body modal frequency is separated from the elastic modal frequency, and the problem of rigid-elastic coupling aeroelastic stability is not easy to occur.
2. The method for improving the pneumatic servo elastic stability of claim 1, wherein the airframe model is linearized at each design dynamic pressure point, the cut-off frequency, the phase angle margin, the crossing frequency and the amplitude margin are calculated by loop-splitting simulation, the robustness of the airframe model is analyzed, and whether the control law dynamic tuning parameters in the full envelope range have consistent control quality is verified, so that the control law robustness and the flight safety of the airplane are improved.
3. The method for improving aeroelastic servo stability of claim 2, wherein, by aeroelastic servo stability analysis, checking whether the amplitude margin of the longitudinal, transverse and course channels of the airplane in the full envelope range meets the condition of more than 6db, and whether the phase margin meets the condition of more than 45 degrees; if not, the corresponding parameter adjustment is carried out through the flight control rhythm dynamic parameter adjustment technology.
4. The method for improving the aerodynamic servo elastic stability as claimed in claim 1, wherein the optimization of the model needs to fully consider an aircraft structure finite element model, so a full-aircraft ground resonance test under multi-point sinusoidal excitation is adopted to obtain data of the model in three aspects of generalized mass, generalized stiffness and structural damping coefficient, the data is used for correcting the full-aircraft finite element model, the modal, slope and control plane inertial coupling parameters of the position of the sensor are corrected through an open loop frequency response characteristic test, and then whether the control law open loop amplitude-frequency curves of a longitudinal channel, a transverse channel and a heading channel after the aircraft correction are close to the test values or not is compared.
5. The method for improving the aerodynamic servo elastic stability of claim 4, wherein a full-aircraft ground structure modal coupling test is performed, the measured open-loop frequency response characteristic test characteristic is used for correcting the modal, slope and control plane inertial coupling parameters of the position where the sensor is located, and then whether the control law open-loop amplitude-frequency curves of the longitudinal channel, the transverse channel and the heading channel after the aircraft is corrected are compared to be close to the test value.
6. The method of claim 4, wherein the body structure rigid natural frequency characteristics and body elastic frequency characteristics of the aircraft are analyzed to avoid coupling of rigid and elastic modes.
7. The method of claim 1, wherein the airplane link transfer function of the airframe model is a ratio of a signal received by a sensor to a control plane deflection.
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| CN112733259B (en) * | 2020-12-29 | 2023-11-21 | 中国航空工业集团公司西安飞机设计研究所 | Pneumatic servo elasticity rapid iterative analysis and design method |
| CN112859592B (en) * | 2020-12-29 | 2022-08-09 | 中国航空工业集团公司西安飞机设计研究所 | Method for controlling structural modal frequency of turboprop aircraft |
| CN113848963B (en) * | 2021-11-29 | 2023-11-28 | 中国航空工业集团公司沈阳飞机设计研究所 | Control law parameter design method of flight control system |
| CN114942652B (en) * | 2022-06-20 | 2024-07-12 | 成都飞机工业（集团）有限责任公司 | Method for calculating phase stability margin of unmanned aerial vehicle flight control system |
| CN116119024B (en) * | 2023-04-17 | 2023-07-18 | 成都沃飞天驭科技有限公司 | Aircraft test platform and design method thereof |
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