CN109733592B - Automatic airplane balancing control method and system - Google Patents
Automatic airplane balancing control method and system Download PDFInfo
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Abstract
The application provides an automatic airplane trim control method and system, wherein the method comprises the following steps: acquiring the pitching rate, the rolling rate, the pedal displacement, the rolling angle, the pitching angle, the attack angle, the time when the longitudinal rod is in the neutral position and the time when the transverse rod is in the neutral position of the airplane; determining whether to execute a pitch angle control rule according to the time when the longitudinal rod is in the neutral position, the pitch rate, the roll angle and the pitch angle; determining whether to execute a roll angle control rule according to the time when the cross rod is in the neutral position, the roll angle rate, the roll angle, the pitch angle and the displacement of the pedals; determining whether to execute a horizontal flight control law according to the time when the longitudinal rod is in the neutral position, the time when the transverse rod is in the neutral position, the pedal displacement and the attack angle; and determining whether to execute an integral control law according to the roll rate, the pedal displacement and the attack angle.
Description
Technical Field
The application relates to the technical field of flight control, and particularly provides an automatic airplane trim control method and system.
Background
Because the trim control surfaces of different flight states of the airplane are different and are continuously disturbed by airflow in the air, a pilot needs to frequently adjust the displacement of a steering column through mechanical adjustment to achieve the purpose of trimming the airplane, and the balancing burden of the pilot is heavy.
Disclosure of Invention
In order to solve at least one of the above technical problems, the present application provides an automatic trim control method and system for an aircraft.
In a first aspect, the present application provides a method for controlling automatic trim of an aircraft, including: acquiring the pitching rate, the rolling rate, the pedal displacement, the rolling angle, the pitching angle, the attack angle, the time when the longitudinal rod is in the neutral position and the time when the transverse rod is in the neutral position of the airplane; determining whether to execute a pitch angle control law according to the time when the longitudinal rod is in the neutral position, the pitch rate, the roll angle and the pitch angle; determining whether to execute a roll angle control rule according to the time when the cross rod is in the neutral position, the roll angle rate, the roll angle, the pitch angle and the displacement of the pedal; determining whether to execute a horizontal flight control law according to the time when the longitudinal rod is in the neutral position, the time when the transverse rod is in the neutral position, the pedal displacement and the attack angle; and determining whether to execute an integral control law according to the roll rate, the pedal displacement and the attack angle.
According to at least one embodiment of the present application, determining whether to execute a pitch angle control law according to the time the pitch stick is in neutral, the pitch rate, the roll angle, and the pitch angle includes: if the time that the longitudinal rod is in the neutral position is equal to the preset time, the pitch rate is smaller than or equal to a first threshold value, the roll angle is smaller than or equal to the preset angle, and the pitch angle is smaller than or equal to the preset angle, executing a pitch angle control law; otherwise, the pitch angle control rule is not executed.
According to at least one embodiment of the present application, the pitch control law is represented by:
wherein N isycmdFor overload instruction, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzIn order to be the pitch angle rate signal,
the pitch angle that needs to be maintained.
According to at least one embodiment of the present application, determining whether to execute a roll angle control law according to a time that the crossbar is in neutral, the roll angle rate, the roll angle, the pitch angle, and the pedal displacement includes: if the time that the cross rod is in the neutral position is equal to the preset time, the roll angle rate is smaller than or equal to a second threshold value, the roll angle is smaller than or equal to the preset angle, the pitch angle is smaller than or equal to the preset angle, and the pedal displacement is smaller than or equal to the preset displacement, executing a roll angle control rule; otherwise, the roll angle control law is not executed.
According to at least one embodiment of the present application, the roll angle control law is represented by:
wherein D isxcmdIn order to be a roll angle command signal,
is a low-pass filter, KΔyIs a roll angle gain coefficient, gammaAt presentAt the current roll angle, KωxAs roll rate gain factor, ωxAs roll rate signal, gammaLocking inThe roll angle that needs to be maintained.
According to at least one embodiment of the present application, determining whether to execute a level flight control law according to the time when the vertical pole is in neutral, the time when the lateral pole is in neutral, the pedal displacement, and the angle of attack comprises: if the time that the longitudinal rod is in the neutral position is longer than the preset time, the time that the transverse rod is in the neutral position is longer than the preset time, the pedal displacement is smaller than or equal to the preset displacement, and the attack angle is smaller than or equal to the preset angle, executing a horizontal flight control law; otherwise, the horizontal flight control law is not executed.
According to at least one embodiment of the present application, the horizontal flight control law is represented by the following formula:
wherein N isycmdIn order to be an overload instruction,
is a low-pass filter, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzFor pitch angle rate signals, alphaFeedbackIs the current angle of attack signal.
According to at least one embodiment of the present application, determining whether to execute an integral control law based on the roll rate, the pedal displacement, and the angle of attack comprises: if the roll rate is smaller than or equal to a second threshold value, the pedal displacement is smaller than or equal to a preset displacement, and the attack angle is smaller than or equal to a preset angle, executing an integral control law; otherwise, the integral control law is not executed.
According to at least one embodiment of the present application, the integral control law is represented by:
wherein D isxcmdAs a roll command signal, DxFor steering column command signals, KDxFor steering column command gain factor, omegaxAs roll rate signal, KωxIs the roll angular rate gain coefficient, vbIn order to obtain the flight speed of the airplane,
as an integrator element, KIIs the integrator gain factor.
In a second aspect, the present application provides an aircraft auto-trim control system, the system comprising a controller configured to execute the aircraft auto-trim control method of the first aspect.
The automatic airplane trim control method and the automatic airplane trim control system have the advantages of being strong in applicability, strong in robustness, convenient to use and the like, and can be applied to most airplane control systems using digital telex only by adaptively changing part of parameters.
Compared with the existing trim controller, the trim controller has the following advantages: mechanical hardware is not needed, the cost of the airplane is reduced, the problems of more additional mechanisms, heavy weight, large occupied space, high failure rate and high maintenance cost of mechanical adjustment are solved, the weight of the airplane is reduced, and the performance of the airplane is improved; the influence of wind disturbance, external hanging asymmetry and manufacturing error can be automatically eliminated, the problem of balancing burden of a pilot is fundamentally solved, and the pilot does not need to carry out continuous fine adjustment according to the flight state and external disturbance.
Drawings
FIG. 1 is a block diagram of an aircraft auto-trim control system provided by an embodiment of the present application;
FIG. 2 is a schematic diagram of a controller control logic provided by an embodiment of the present application;
FIG. 3 is a schematic diagram of another controller control logic provided by an embodiment of the present application;
FIG. 4 is a schematic diagram of another controller control logic provided by an embodiment of the present application;
FIG. 5 is a schematic diagram of a controller control logic according to an embodiment of the present application;
FIG. 6 is a schematic flow chart diagram illustrating a method for controlling automatic trim of an aircraft according to an embodiment of the present disclosure;
fig. 7 is a structural diagram of a pitch angle control law provided in an embodiment of the present application;
FIG. 8 is a block diagram illustrating roll angle control rules provided in an embodiment of the present application;
FIG. 9 is a block diagram of the horizontal flight control law provided by the embodiments of the present application;
fig. 10 is a structural diagram of an integral control law provided in an embodiment of the present application;
Detailed Description
The present application will be described in further detail with reference to the following drawings and examples. It is to be understood that the specific embodiments described herein are merely illustrative of the relevant application and are not limiting of the application. It should be noted that, for convenience of description, only the portions related to the present application are shown in the drawings.
It should be noted that the embodiments and features of the embodiments in the present application may be combined with each other without conflict. The present application will be described in detail below with reference to the embodiments with reference to the attached drawings.
Fig. 1 is a block diagram of an aircraft automatic trimming control system provided in an embodiment of the present application, and as shown in fig. 1, the aircraft automatic trimming control system includes a controller 101, a pitch angle holder 102, a roll angle holder 103, a horizontal flight controller 104, and an integrator 105, where the pitch angle holder 102, the roll angle holder 103, the horizontal flight controller 104, and the integrator 105 are all connected to the controller 101.
The controller 101 can control whether or not the pitch angle holder 102, the roll angle holder 103, the horizontal flight controller 104, and the integrator 105 are operated, and for example, the controller 101 can transmit a high/low level to the pitch angle holder 102, the roll angle holder 103, the horizontal flight controller 104, and the integrator 105, and start operating when receiving a high level, and does not operate when receiving a low level.
It should be noted that the controller 101 is intended to maintain compliance with the pilot's maneuver "intent", to automatically complete the trim maneuver when the pilot "desires", and to exit in time with no (or very little) transients when the pilot needs to maneuver the aircraft.
The most important principle of the design of the controller 101 is that it cannot be violated by the pilot's manoeuvre, i.e. contrary to the pilot's control commands. For example, when the pilot is performing a lateral maneuver, the auto-trim controller cannot output a crossbar command as opposed to the pilot's maneuver; when the airplane is asymmetrically hung, the intelligent controller can automatically identify the asymmetric characteristic of the airplane, so that the force of a steering column is eliminated. In addition, the on-off and fault transients of the controller 101 should be small and even hard to detect by the pilot.
In one example, as shown in fig. 2, when the controller 101 receives the auto-trim function on signal, and the time when the pitch stick of the aircraft is in the neutral state is equal to the preset time, and the pitch rate is less than or equal to the first threshold, and the roll angle is less than or equal to the preset angle, and the pitch angle is less than or equal to the preset angle, the controller 101 sends a high level to the pitch angle holder 102, so that the pitch angle holder 102 is in the working state; if at least one of the above conditions is not met, the controller 101 sends a low level to the pitch angle holder 102, causing the pitch angle holder 102 to be inoperative.
In one example, as shown in fig. 3, when the controller 101 receives the auto-trim function on signal, while the crossbar of the aircraft is in the neutral state for a preset time, the roll angle rate is less than or equal to the second threshold, the roll angle is less than or equal to the preset angle, the pitch angle is less than or equal to the preset angle, and the pedal displacement is less than or equal to the preset displacement, the controller 101 sends a high level to the roll angle holder 103, so that the roll angle holder 103 is in the working state; if at least one of the above conditions is not met, the controller 101 may send a low level to the roll angle keeper 103 to disable the roll angle keeper 103.
In an example, as shown in fig. 4, when the controller 101 receives an auto-trim function on signal, and at the same time, the time that the vertical bar of the aircraft is in the neutral state is longer than the preset time, the time that the horizontal bar is in the neutral state is longer than the preset time, the pedal displacement is less than or equal to the preset displacement, and the attack angle is less than or equal to the preset angle, the controller 101 sends a high level to the horizontal flight controller 104, so that the horizontal flight controller 104 is in an operating state; if at least one of the above conditions is not satisfied, the controller 101 sends a low level to the horizontal flight controller 104 to disable the horizontal flight controller 104.
In an example, as shown in fig. 5, when the controller 101 receives the auto-trim function start signal, and at the same time, the roll rate of the aircraft is less than or equal to the second threshold, the pedal displacement is less than or equal to the preset displacement, and the attack angle is less than or equal to the preset angle, the controller 101 sends a high level to the integrator 105, so that the integrator 105 is in an operating state; if at least one of the above conditions is not met, the controller 101 sends a low level to the integrator 105 to disable the integrator 105.
Fig. 6 is a schematic flow chart of an aircraft automatic trim control method provided in an embodiment of the present application, and as shown in fig. 6, the method includes the following steps:
step 601, acquiring the pitching rate, the rolling rate, the pedal displacement, the rolling angle, the pitching angle, the attack angle, the time when the longitudinal rod is in the neutral position and the time when the cross rod is in the neutral position of the airplane.
In the present embodiment, the controller 101 acquires the pitch rate, roll rate, pedal displacement, roll angle, pitch angle, attack angle, time when the pitch beam is in neutral, and time when the lateral beam is in neutral of the aircraft in real time.
In one example, the pitch rate, roll rate, foot pedal displacement, roll angle, pitch angle, angle of attack, time the pitch stick is neutral, and time the crossbar is neutral of the aircraft may all be acquired by sensors on the aircraft.
When the controller 101 receives the start signal of the automatic balancing function, the corresponding control law can be selected according to the pitch rate, the roll rate, the pedal displacement, the roll angle, the pitch angle, the attack angle, the time when the vertical rod is in the neutral position, and the time when the horizontal rod is in the neutral position to perform the balancing of the aircraft.
Step 602, determining whether to execute a pitch angle control law according to the time when the pitch stick is in the neutral position, the pitch rate, the roll angle and the pitch angle.
In some embodiments, if the time that the pitch stick is in the neutral position is equal to a preset time, and the pitch rate is equal to or less than a first threshold, and the roll angle is equal to or less than a preset angle, and the pitch angle is equal to or less than a preset angle, then executing a pitch angle control law; otherwise, the pitch angle control law is not executed.
The pitch angle control law is a typical longitudinal trim mode in flight, the main flight stages of the pitch angle control law comprise that an accelerating pitch angle after the aircraft flies off and off keeps climbing, a pitch angle when the aircraft climbs to a cruising altitude or dives to a landing route keeps climbing/gliding, a pitch angle when the ground/sea is attacked keeps diving attack and the like, the essence is that a controller executes a specified control mode when judging that an expected trim state is the pitch angle maintenance, the structural diagram is shown in fig. 7, and a simplified control instruction forming calculation formula is as follows:
wherein N isycmdFor overload instruction, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzIn order for the pitch angle rate signal to be,
the pitch angle that needs to be maintained.
Step 603, determining whether to execute a roll angle control rule according to the time when the cross rod is in the neutral position, the roll angle rate, the roll angle, the pitch angle and the displacement of the pedal.
In some embodiments, if the time that the crossbar is in the neutral is equal to the preset time, the roll rate is less than or equal to the second threshold, the roll angle is less than or equal to the preset angle, the pitch angle is less than or equal to the preset angle, and the pedal displacement is less than or equal to the preset displacement, then executing a roll angle control law; otherwise, the roll angle control law is not executed.
Roll control laws are the typical lateral trim approach in flight, and in most steady-state flights, a fixed roll angle, such as hover, needs to be maintained. The zero roll angle can be considered to be maintained in the flight state without the roll angle, such as horizontal flight, and the control law design principle is similar to the roll angle maintenance. The essential controller executes a designated control mode when judging that the expected trim state is roll angle maintenance, and the structure diagram is shown in fig. 8, and the simplified control command forming calculation formula is as follows:
wherein D isxcmdIn order to be a roll angle command signal,
is a low-pass filter, KΔyIs a roll angleGain factor, gammaAt presentAt the current roll angle, KωxAs roll rate gain factor, ωxAs roll rate signal, gammaLocking inThe roll angle that needs to be maintained.
And step 604, determining whether to execute a horizontal flight control law according to the time when the longitudinal rod is in the neutral position, the time when the transverse rod is in the neutral position, the pedal displacement and the attack angle.
In some embodiments, if the time that the longitudinal rod is in the neutral position is longer than the preset time, the time that the transverse rod is in the neutral position is longer than the preset time, the pedal displacement is less than or equal to the preset displacement, and the attack angle is less than or equal to the preset angle, the horizontal flight control law is executed; otherwise, the horizontal flight control law is not executed.
The horizontal flight control law is a typical longitudinal trim mode in flight, and the applied flight phase comprises all horizontal flight states, such as cruise, flight routes, refueling and the like. The essence is that the controller executes a designated control mode when judging that the desired trim state is a height hold, and the structure diagram is as follows as the simplified control command forming calculation formula shown in fig. 9:
wherein N isycmdIn order to be an overload instruction,
is a low-pass filter, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzFor pitch angle rate signals, alphaFeedbackIs the current angle of attack signal.
Step 605, determining whether to execute an integral control law according to the roll rate, the pedal displacement and the attack angle.
In some embodiments, if the roll rate is less than or equal to the second threshold, the pedal displacement is less than or equal to the preset displacement, and the attack angle is less than or equal to the preset angle, executing an integral control law; otherwise, the integral control law is not executed.
The integral control law is used for transverse horizontal balancing of the airplane and mainly overcomes asymmetric moment caused by asymmetric plug-in of the airplane by generating a transverse steering rod instruction. Since most aircraft are not provided with integrators in the transverse direction at present, in order to realize the non-static control, the integrators must be added.
The integrator is effective for aircraft with asymmetric plug-ins or other sustained roll torques, which greatly reduces the pilot's steering burden. The essence of the method is that the intelligent decision controller executes a designated control mode when judging that the aircraft has asymmetric rolling moment, the structure diagram is shown in fig. 10, and the simplified control instruction forming calculation formula is as follows:
wherein D isxcmdAs a roll command signal, DxFor steering column command signals, KDxFor steering column command gain factor, omegaxAs roll rate signal, KωxIs the roll angular rate gain coefficient, vbIn order to obtain the flight speed of the airplane,
as an integrator element, KIIs the integrator gain factor.
It should be noted that the preset time, the first threshold, the second threshold and the preset angle in the above embodiments may be flexibly set by those skilled in the art according to actual control needs, and are not limited herein.
In a second aspect, the present application further provides an automatic aircraft trim control system, which includes a controller, and the controller is used in the automatic aircraft trim control method in the above embodiment.
So far, the technical solutions of the present application have been described in connection with the preferred embodiments shown in the drawings, but it is easily understood by those skilled in the art that the scope of the present application is obviously not limited to these specific embodiments. Equivalent changes or substitutions of related technical features can be made by those skilled in the art without departing from the principle of the present application, and the technical scheme after the changes or substitutions will fall into the protection scope of the present application.
Claims (10)
1. An aircraft auto-trim control method, comprising:
acquiring the pitching rate, the rolling rate, the pedal displacement, the rolling angle, the pitching angle, the attack angle, the time when the longitudinal rod is in the neutral position and the time when the transverse rod is in the neutral position of the airplane;
determining whether to execute a pitch angle control law according to the time when the longitudinal rod is in the neutral position, the pitch rate, the roll angle and the pitch angle;
determining whether to execute a roll angle control rule according to the time when the cross rod is in the neutral position, the roll angle rate, the roll angle, the pitch angle and the displacement of the pedal;
determining whether to execute a horizontal flight control law according to the time when the longitudinal rod is in the neutral position, the time when the transverse rod is in the neutral position, the pedal displacement and the attack angle;
and determining whether to execute an integral control law according to the roll rate, the pedal displacement and the attack angle.
2. The aircraft auto-trim control method of claim 1, wherein determining whether to execute a pitch angle control law as a function of the time the pitch bar is neutral, the pitch rate, the roll angle, and the pitch angle comprises:
if the time that the longitudinal rod is in the neutral position is equal to the preset time, the pitch rate is smaller than or equal to a first threshold value, the roll angle is smaller than or equal to the preset angle, and the pitch angle is smaller than or equal to the preset angle, executing a pitch angle control law; otherwise, the pitch angle control rule is not executed.
3. The method of claim 2, wherein the pitch control law is expressed by the following equation:
wherein N isycmdFor overload instruction, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzIn order to be the pitch angle rate signal,
the pitch angle that needs to be maintained.
4. The aircraft auto-trim control method of claim 1, wherein determining whether to execute a roll angle control law based on the time the crossbar is in neutral, the roll rate, the roll angle, the pitch angle, and the pedal displacement comprises:
if the time that the cross rod is in the neutral position is equal to the preset time, the rolling angle rate is smaller than or equal to a second threshold value, the rolling angle is smaller than or equal to the preset angle, the pitch angle is smaller than or equal to the preset angle, and the pedal displacement is smaller than or equal to the preset displacement, executing a rolling angle control rule; otherwise, the roll angle control law is not executed.
5. The aircraft auto-trim control method of claim 4, wherein the roll angle control law is expressed by:
wherein D isxcmdIn order to be a roll angle command signal,
is the transfer function of a low-pass filter, KΔyIs a roll angle gain coefficient, gammaAt presentAt the current roll angle, KωxAs roll rate gain factor, ωxAs roll rate signal, gammaLocking inThe roll angle that needs to be maintained.
6. The aircraft auto-trim control method of claim 1, wherein determining whether to execute a level flight control law based on the time the longitudinal bar is neutral, the time the lateral bar is neutral, the foot pedal displacement, and the angle of attack comprises:
if the time that the longitudinal rod is in the neutral position is longer than the preset time, the time that the transverse rod is in the neutral position is longer than the preset time, the pedal displacement is smaller than or equal to the preset displacement, and the attack angle is smaller than or equal to the preset angle, executing a horizontal flight control law; otherwise, the horizontal flight control law is not executed.
7. The aircraft auto-trim control method of claim 6, wherein the horizontal flight control law is represented by:
wherein N isycmdIn order to be an overload instruction,
is the transfer function of a low-pass filter, KΔθIn order to be the gain factor,
for the current pitch angle signal, Kωz-αFor pitch angle rate gain, ωzFor pitch angle rate signals, alphaFeedbackIs the current angle of attack signal.
8. The aircraft auto-trim control method of claim 1, wherein determining whether to execute an integral control law based on the roll rate, the pedal displacement, and the angle of attack comprises:
if the roll rate is smaller than or equal to a second threshold value, the pedal displacement is smaller than or equal to a preset displacement, and the attack angle is smaller than or equal to a preset angle, executing an integral control law; otherwise, the integral control law is not executed.
9. The aircraft auto-trim control method of claim 8, wherein the integral control law is expressed by:
wherein D isxcmdAs a roll command signal, DxFor steering column command signals, KDxFor steering column command gain factor, omegaxAs roll rate signal, KωxIs the roll angular rate gain coefficient, vbIn order to obtain the flight speed of the airplane,
as an integrator element, KIIs the integrator gain factor.
10. An aircraft auto-trim control system, characterized in that the system comprises a controller for performing the aircraft auto-trim control method of any of claims 1 to 9.
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| CN110989680B (en) * | 2019-12-25 | 2024-02-02 | 中国航空工业集团公司沈阳飞机设计研究所 | Flight control guiding method |
| TWI736139B (en) * | 2020-02-17 | 2021-08-11 | 經緯航太科技股份有限公司 | UAV with night flight operation assistance function |
| CN113460289B (en) * | 2021-09-06 | 2022-04-22 | 中国商用飞机有限责任公司 | Method, system and device for pitch trim feedback of aircraft |
| CN113479316B (en) * | 2021-09-08 | 2021-11-30 | 中国商用飞机有限责任公司 | Methods, systems, and media for controlling horizontal tail unloading of an aircraft |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3258517A (en) * | 1961-08-02 | 1966-06-28 | Communications Patents Ltd | Ground based flight training apparatus |
| US4220993A (en) * | 1978-09-15 | 1980-09-02 | Sperry Corporation | Engine throttle control system |
| US4392203A (en) * | 1981-03-30 | 1983-07-05 | United Technologies Corporation | Aircraft coordinated turn with lagged roll rate |
| US5001646A (en) * | 1988-12-19 | 1991-03-19 | Mcdonnell Douglas Corporation | Automated helicopter flight control system |
| CN101287647A (en) * | 2005-08-31 | 2008-10-15 | 贝尔直升机特克斯特龙有限公司 | Method and apparatus for improving main rotor yoke fatigue life |
| CN203444337U (en) * | 2013-08-27 | 2014-02-19 | 长春富维—江森自控汽车饰件系统有限公司 | Experiment angle adjusting device for automobile interior decorative part |
| WO2017073726A1 (en) * | 2015-10-28 | 2017-05-04 | 旭硝子株式会社 | Stain-proof article and production method for stain-proof article |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| BR112015032444A2 (en) * | 2013-06-28 | 2017-07-25 | Ge Aviat Systems Ltd | method for diagnosing a horizontal stabilizer system failure in an aircraft |
-
2018
- 2018-11-23 CN CN201811409844.6A patent/CN109733592B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3258517A (en) * | 1961-08-02 | 1966-06-28 | Communications Patents Ltd | Ground based flight training apparatus |
| US4220993A (en) * | 1978-09-15 | 1980-09-02 | Sperry Corporation | Engine throttle control system |
| US4392203A (en) * | 1981-03-30 | 1983-07-05 | United Technologies Corporation | Aircraft coordinated turn with lagged roll rate |
| US5001646A (en) * | 1988-12-19 | 1991-03-19 | Mcdonnell Douglas Corporation | Automated helicopter flight control system |
| CN101287647A (en) * | 2005-08-31 | 2008-10-15 | 贝尔直升机特克斯特龙有限公司 | Method and apparatus for improving main rotor yoke fatigue life |
| CN203444337U (en) * | 2013-08-27 | 2014-02-19 | 长春富维—江森自控汽车饰件系统有限公司 | Experiment angle adjusting device for automobile interior decorative part |
| WO2017073726A1 (en) * | 2015-10-28 | 2017-05-04 | 旭硝子株式会社 | Stain-proof article and production method for stain-proof article |
Non-Patent Citations (2)
| Title |
|---|
| Multiple time scale analysis of aircraft longitudinal dynamics with aerodynamic vectoring;Adnan Maqsood，Tiauw Hiong Go;《Nonlinear Dynamics》;20111222;731-742 * |
| 民用飞机飞行仿真系统研究;曹全新,严林芳,刘永超;《测控技术》;20090718;102-107 * |
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