CN111190437B - Control method and flight control system for rolling torque under large attack angle - Google Patents
Control method and flight control system for rolling torque under large attack angle Download PDFInfo
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- CN111190437B CN111190437B CN202010000819.3A CN202010000819A CN111190437B CN 111190437 B CN111190437 B CN 111190437B CN 202010000819 A CN202010000819 A CN 202010000819A CN 111190437 B CN111190437 B CN 111190437B
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
- B64C23/06—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/10—Drag reduction
Abstract
The invention provides a control method of rolling torque under a large attack angle and a flight control system. The control method controls the roll torque by taking the duty ratio of the jet flow as a control quantity, and controls the duty ratio of the jet flow through the high-frequency electromagnetic valve under the condition of ensuring constant supply air pressure, so that the speed of the jet flow in the air outlet time period is kept constant, and the speed of the jet flow in the air closing time period is zero, and the control effect without a steering effect dead zone is realized. The jet flow exciter is connected with the flight control system, the traditional control surface is replaced, the defect that the traditional control surface fails under the condition of large attack angle flow separation is overcome, and the flight envelope of the aircraft is further widened.
Description
Technical Field
The invention discloses a rolling torque control method under a large attack angle and a flight control system, and belongs to the field of aircraft control.
Background
At high angles of attack, the conventional roll control surface (aileron) fails due to severe boundary layer separation, and it becomes very inefficient and dangerous to use it to control over-stall maneuvers at high angles of attack. There is a need to design an effective and linear control method to control the roll torque of an aircraft at a large angle of attack.
The surface of the aircraft with the swept wing can wind two or more vortex systems from the front edge, and because the vortex systems become important components of the lift force under a large attack angle, the stress condition of the aircraft can be effectively controlled if the vortex systems can be controlled.
Jet flow along the leading edge is a way of controlling the vortex system at high angles of attack, the main principle being to inject momentum into the vortex core and thereby delay the breaking of the vortex at high angles of attack, but if this control is used to control the roll torque of the aircraft there is a disadvantage in that the control is ineffective, i.e. there is a "dead zone" in the control, when the velocity of the jet flow along the leading edge does not reach a certain threshold. To illustrate this, fig. 1 shows the variation process of the roll moment with the jet flow momentum coefficient at 26 degrees of attack angle for a certain type of flying wing type configuration aircraft, and it can be seen that if the roll moment is controlled by controlling the jet flow momentum coefficient, there will be a very obvious "steering effect dead zone". How to design a roll torque control mode which has no dead zone and has a linear control effect is a problem to be solved.
Disclosure of Invention
The purpose of the invention is as follows: aiming at the problems in the prior art, the invention provides a control method for controlling the roll torque of an aircraft by controlling the jet flow duty ratio, so that the control effect without a steering effect dead zone and approximate linearity is realized.
Another object of the present invention is to provide a flight control system implementing the above control method.
The technical scheme is as follows: according to the first aspect of the invention, the method for controlling the rolling torque under the large attack angle is applied to a swept wing aircraft, the method utilizes pulse jet flow to control a vortex system generated in flight, and controls the rolling torque of the aircraft by changing the duty ratio of the pulse jet flow, wherein the increment of the rolling torque generated by blowing is in a proportional relation with the duty ratio of the jet flow, the proportional relation means that a larger duty ratio is set if large rolling torque is needed, and a smaller duty ratio is set if small rolling torque is needed, and the increment of the rolling torque generated by blowing and the duty ratio of the jet flow are basically in a linear relation.
Further, the switch of the pulse jet is controlled by high and low level signals, wherein the high level signal is used for opening the jet, and the low level signal is used for closing the jet.
Further, the high-low level signal is obtained by the following method: the flight control command and the sensor data are transmitted into a flight control system of the aircraft, corresponding steering engine signals are obtained through calculation of a flight control algorithm, and then the steering engine signals are converted into high and low level signals with required duty ratios to serve as control signals of the pulse jet.
According to a second aspect of the present invention, there is provided a flight control system for implementing the above control method, comprising:
the flight control panel is used for calculating a steering engine signal according to flight task requirements;
the signal conversion module is used for converting a traditional steering engine signal output by the flight control panel into a square wave signal with a required duty ratio, and the duty ratio is set according to the magnitude of the required rolling torque and basically linearly changes along with the magnitude of the rolling torque;
and the jet flow exciter is used for emitting pulse type jet flow under the control of the square wave signal output by the signal conversion module, wherein the high-level time jet flow of the square wave signal is opened, and the low-level time jet flow of the square wave signal is closed.
Has the advantages that:
1. the invention controls the duty ratio of the jet flow through the high-frequency electromagnetic valve under the condition of ensuring that the supply air pressure is constant, so that the speed of the jet flow is kept constant in the air outlet time period and is zero in the air closing time period. This also promotes the control effect of "must be effective" to give vent to anger to realize the control effect of no steering effect blind spot.
2. The control principle of the jet flow control mode is similar to the working principle of adjusting the brightness of the LED, and the linear control of the rolling torque can be realized by skillfully changing the duty ratio of the square wave signal.
3. The jet flow exciter is connected with the flight control system, the traditional control surface is replaced, the defect that the traditional control surface fails under the condition of large attack angle flow separation is overcome, and the flight envelope of the aircraft is further widened.
Drawings
FIG. 1 is a graph showing the variation process of roll torque with the momentum coefficient of jet flow at an attack angle of 26 degrees;
FIG. 2 is a schematic view of the location of a jet actuator along a leading edge in accordance with the present invention;
FIG. 3 illustrates the control effect of a pulsed jet along the leading edge at an angle of attack of 26 degrees in accordance with the present invention;
FIG. 4 is a schematic diagram of a duty cycle signal acquisition process according to the present invention;
fig. 5 is a block diagram of a control system according to the present invention.
Detailed Description
The technical scheme of the invention is further explained by combining the attached drawings.
The invention provides a method for replacing a traditional aileron by utilizing jet flow along the leading edge under a large attack angle, which is different from the traditional idea that the blowing amount (namely the jet flow momentum coefficient) is taken as the control amount, the control amount is the duty ratio of the jet flow along the leading edge, and the effective and linear rolling torque control effect can be provided under the large attack angle. Jet duty cycle refers to the ratio of the time a jet is fired during a cycle to the entire cycle.
An effective edge jet flow exciter is designed aiming at the aerodynamic configuration of an aircraft, and aims to control the front edge vortex shaft adverse pressure gradient of a backswept wing aircraft, delay vortex breaking, enable the local lift force of a blowing side to rise and generate corresponding rolling torque. Typical jet actuators applied on certain types of flying wing aircraft are arranged along the leading edge on both sides of the aircraft head, as shown in fig. 2 at 1, with the jet direction being aft in the direction of the leading edge. In operation, a pulsed jet is emitted from the leading edge, thereby affecting the shape of the leading edge vortex. When the pulse type jet flow exciter works, the air pressure value at the inlet of the pulse type jet flow exciter is constant, so that the transient speed of jet flow is almost constant; the working frequency of the exciter is also kept unchanged; during control, it is the duty cycle of the jet that is varied.
Aiming at the flying wing layout aircraft, the invention carries out a plurality of tests aiming at pulsed jet flow along the leading edge at different attack angles, fig. 3 shows the control effect of the pulsed jet flow along the leading edge under the attack angle of 26 degrees, a curve graph of the rolling torque coefficient changing along with the change of the duty ratio of the pulsed jet flow at the right side is shown in the graph, and the working frequency of the jet flow in the test is 10 Hz. It can be seen from the figure that the larger the duty ratio is, the larger the rolling torque is, the substantially linear relationship between the duty ratio and the rolling torque is, and the rolling torque can be changed by changing the duty ratio.
In the actual implementation process, as shown in fig. 4, the flight control instruction and the data (such as attitude, position, etc.) of the airborne sensor are transmitted to the flight control system, the corresponding steering engine signal is obtained by calculation through a flight control algorithm, and then the steering engine signal is converted into a control signal with a required duty ratio through a conversion module, so as to control the opening and closing of the actuator. The flight control algorithm may use a conventional PID (proportional-derivative-integral) algorithm, which is not the core of the present invention and will not be described herein. The acquisition of the control signal to the fluidic actuator is described in detail below.
In one embodiment, the general arrangement of the control system is as shown in fig. 5, before control is performed, a high-frequency solenoid valve is connected with a high-pressure gas cylinder, and the gas cylinder is connected with the solenoid valve through a pressure reducing valve, so that the gas cylinder can supply a constant pressure to the solenoid valve. The exciter can be realized through the electromagnetic valve, the air outlet and the air bottle. The flight control panel is fixedly connected with the aircraft and can generate corresponding control signals through a flight control algorithm according to the control instructions. The flight control panel adopts a pixhawk open-source flight control panel, and the traditional PID algorithm is used for control. And the PWM signal output by the flight control panel is used for controlling the on-off of the electromagnetic valve. Because the working voltage of the electromagnetic valve is 24V, and the signal output by the flight control panel is a PWM signal with the pulse width of 1-2ms, the period of 20ms and the amplitude of 5V, in other words, the signal duty ratio range is between 5% -10%, the signal firstly does not meet the requirement of the voltage amplitude, secondly, the duty ratio in this range cannot exert the effect of this control method very delicately, therefore, the invention converts the signal into a square wave signal with the duty ratio of 0-70% and the amplitude of 24V through a signal conversion module according to the pulse width value of the flight control panel signal, and specifically, the signal conversion module comprises a signal conversion panel and a voltage amplification module, the signal conversion board is an STM32 board, the flight control board outputs PWM waves, the STM32 board receives signals, converts the duty ratio to output the PWM waves with the duty ratio of 0-70%, and the voltage is amplified through the voltage amplification module. The formula for calculating the duty ratio of the output signal of the signal conversion plate is shown in the following formula:
in the formula, DC refers to the duty ratio of the output signal of the signal conversion plate, t is the pulse width of the steering engine signal currently output by the flight control plate, and t isminThe minimum value of the pulse width of the signals of the flight control plate steering engine is obtained. The upper limit of the duty ratio is set to 70% in order to protect the high-frequency solenoid valve from damage, and the upper limit may be set according to the actual condition of the solenoid valve.
It should be further noted that the above control method is suitable for flight control of most swept-wing aircraft under a large attack angle, but the swept-wing aircraft also has many types, and some may need to design the position of the air outlet according to practical situations, for example, the air outlet may not be arranged at the head, but the duty ratio control method is still suitable. It will be appreciated by those skilled in the art that modifications and equivalents may be made to the embodiments of the invention described herein and thus are intended to be included within the scope of the claims of this invention without departing from the spirit and scope of the invention.
Claims (3)
1. A control method of rolling torque under a large attack angle is applied to a sweepback wing aircraft, and is characterized in that a pulse jet is utilized to control a vortex system generated in flight, and the rolling torque of the aircraft is controlled by changing the duty ratio of the pulse jet, wherein the increment of the rolling torque generated by blowing is in a proportional relation with the duty ratio of the jet;
the switch of the pulse type jet flow is controlled by a high-low level signal, wherein the high level signal is used for opening the jet flow, and the low level signal is used for closing the jet flow; the high-low level signal is obtained by the following method: the flight control command and the sensor data are transmitted into a flight control system of the aircraft, corresponding steering engine signals are obtained through calculation of a flight control algorithm, and then the steering engine signals are converted into high and low level signals with required duty ratios to serve as control signals of the pulse jet.
2. A flight control system for implementing the method of controlling roll torque at high angles of attack of claim 1, comprising:
the flight control panel is used for calculating a steering engine signal according to flight task requirements;
the signal conversion module is used for converting a traditional steering engine signal output by the flight control panel into a square wave signal with a required duty ratio, and the duty ratio is set according to the magnitude of the required rolling torque and basically linearly changes along with the magnitude of the rolling torque;
and the jet flow exciter is used for emitting pulse type jet flow under the control of the square wave signal output by the signal conversion module, wherein the high-level time jet flow of the square wave signal is opened, and the low-level time jet flow of the square wave signal is closed.
3. The flight control system according to claim 2, wherein the jet actuator comprises an air outlet, an air cylinder and an electromagnetic valve, the air cylinder is connected with the electromagnetic valve through a pressure reducing valve, the electromagnetic valve is connected with the signal conversion module, and the switching is realized under the control of the output signal of the signal conversion module.
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| CN202010000819.3A CN111190437B (en) | 2020-01-02 | 2020-01-02 | Control method and flight control system for rolling torque under large attack angle |
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| CN202010000819.3A CN111190437B (en) | 2020-01-02 | 2020-01-02 | Control method and flight control system for rolling torque under large attack angle |
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| CN113682466A (en) * | 2021-09-30 | 2021-11-23 | 中国人民解放军国防科技大学 | Aircraft non-control surface flight control method based on synthetic double-jet flow field control |
| CN117289712A (en) * | 2023-11-27 | 2023-12-26 | 中国航空研究院 | Virtual control surface jet flow control system and method |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102320375A (en) * | 2011-06-27 | 2012-01-18 | 南京航空航天大学 | Aircraft forebody asymmetric vortex control device and control method thereof |
| CN102390525A (en) * | 2011-10-20 | 2012-03-28 | 南京航空航天大学 | Large attack angle asymmetric vortex/lateral force closed-loop active control device |
| CN102417031A (en) * | 2011-10-20 | 2012-04-18 | 南京航空航天大学 | Unsteady small-perturbation control device for synthetic jet of asymmetric vortices at high attack angle |
| CN102602541A (en) * | 2012-03-20 | 2012-07-25 | 南京航空航天大学 | Method for using plasma exciters to control aircraft attitude |
| CN105864232A (en) * | 2016-03-31 | 2016-08-17 | 南京航空航天大学 | Method for recovering tail spin by controlling fore-body vortex and flow control exciter |
| CN205823803U (en) * | 2016-03-31 | 2016-12-21 | 南京航空航天大学 | A kind of double film two-chamber zero-net-mass-flux jet activator |
| CN108116664A (en) * | 2017-12-20 | 2018-06-05 | 南京航空航天大学 | Adaptive excitation control system based on plasma synthesis jet-flow excitor |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103488814B (en) * | 2013-08-16 | 2017-02-15 | 北京航空航天大学 | Closed loop simulation system suitable for controlling attitude of reentry vehicle |
| CN107037824B (en) * | 2017-06-09 | 2023-10-24 | 中国航空工业集团公司哈尔滨空气动力研究所 | Transverse control device and control method for flying wing model |
-
2020
- 2020-01-02 CN CN202010000819.3A patent/CN111190437B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102320375A (en) * | 2011-06-27 | 2012-01-18 | 南京航空航天大学 | Aircraft forebody asymmetric vortex control device and control method thereof |
| CN102390525A (en) * | 2011-10-20 | 2012-03-28 | 南京航空航天大学 | Large attack angle asymmetric vortex/lateral force closed-loop active control device |
| CN102417031A (en) * | 2011-10-20 | 2012-04-18 | 南京航空航天大学 | Unsteady small-perturbation control device for synthetic jet of asymmetric vortices at high attack angle |
| CN102602541A (en) * | 2012-03-20 | 2012-07-25 | 南京航空航天大学 | Method for using plasma exciters to control aircraft attitude |
| CN105864232A (en) * | 2016-03-31 | 2016-08-17 | 南京航空航天大学 | Method for recovering tail spin by controlling fore-body vortex and flow control exciter |
| CN205823803U (en) * | 2016-03-31 | 2016-12-21 | 南京航空航天大学 | A kind of double film two-chamber zero-net-mass-flux jet activator |
| CN108116664A (en) * | 2017-12-20 | 2018-06-05 | 南京航空航天大学 | Adaptive excitation control system based on plasma synthesis jet-flow excitor |
Non-Patent Citations (3)
| Title |
|---|
| 大迎角前体涡控制方法综述;翟建等;《空气动力学学报》;20170630;第35卷(第3期);第354-367页 * |
| 等离子体射流控制机翼气动力矩的实验研究;刘汝兵等;《工程力学》;20160331;第33卷(第3期);第232-238页 * |
| 飞翼布局飞行器等离子体激励增升技术试验研究;姚军锴等;《空气动力学学报》;20180831;第36卷(第4期);第664-669页 * |
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