CN109592064A - Aircraft and mechanical manoeuvring system uneven deformation influence design method to manoeuvre - Google Patents
Aircraft and mechanical manoeuvring system uneven deformation influence design method to manoeuvre Download PDFInfo
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Abstract
The invention discloses aircraft and mechanical manoeuvring system uneven deformations to influence design method to manoeuvre, comprising the following steps: 1) according to the requirement of when air maneuver resultant force and resultant moment, the rudder face degree of bias needed for longitudinal and transverse, course maneuvering when calculating air maneuver；2) bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver；3) the cockpit manipulation displacement of aircraft and steering force during calculating air maneuver；4) different target overload when according to air maneuver, repeat step 1- step 3, obtain longitudinal and transverse, course cockpit manipulation displacement when air maneuver, rod gradient, the bar displacement gradient that steering force overloads corresponding target, the present invention realizes aircraft, and awing true mechanical manoeuvring system characteristic determines process, and true air maneuver handling characteristic designs the method for determination, it ensure that the consistency of flight and design, have modified mechanical manoeuvring system characteristic error, the safety for improving aircraft flight, improves handling quality.
Description
Technical field
The present invention relates to aircraft controllability stability Design technical field more particularly to aircrafts and mechanical manoeuvring system to deform
Difference influences design method to manoeuvre.
Background technique
Currently, although using the steerable system of electrical transmission and optical transport in modern aircraft using more and more extensive,
It is the mainstream in the aircraft or current service aircraft using mechanical system manipulation.
Fighter plane, the bomber of mechanical system manipulation belong to the technology of second-generation aircraft, meter Ge -21, rice such as the former Soviet Union
Lattice -23 and China destroy -7, destroy -8 and be all made of hydraulic booster machinery steerable system, although having become main force's services at present
The third generation or even Fourth Generation Fighters, bomber largely introduce fly-by-wire, but driving mechanism in fly-by-wire
(steering engine) is to can also cause relative deformation when arranging unreasonable between enlarger (booster), and mechanical system manipulation flies
Machine be still be active service operational aircraft and a large amount of mainstream civilian, in general-purpose aircraft.
It is not mentioned also in the description designed both at home and abroad air maneuver handling characteristic at present and considers that airframe is with mechanical
The influence of system phase enantiotropy, although many aircrafts are the absence of systematic true according to the system for modified result aircraft of taking a flight test
Setting analysis method.
Consider and determine that aircraft mechanical manoeuvring system and aircraft deformation are asynchronous right the invention proposes a kind of comprehensive, system
The method that air maneuver handling characteristic influences, so as to ensure that the full mission profile of aircraft meets air maneuver handling characteristic and set
Meter requires, and prevention and prevention aircraft occur non-instruction sexual deviation when friction speed is motor-driven, guarantees flight safety, mitigates driver
Burden, and improve handling quality, to improve safety and the comfort of aircraft.
Summary of the invention
The purpose of the present invention: propose that a kind of aircraft and mechanical manoeuvring system uneven deformation influence design side to manoeuvre
Method can meet air maneuver handling characteristic design requirement ensuring the full mission profile of aircraft, and prevention and prevention aircraft are in difference
Occur non-instruction sexual deviation when speed maneuver, guarantees flight safety, mitigates driver's burden, and improve handling quality, to mention
The safety of high aircraft and comfort.
Technical solution of the present invention:
Aircraft and mechanical manoeuvring system uneven deformation influence design method to manoeuvre, comprising the following steps:
Step 1: according to the requirement of resultant force and resultant moment when air maneuver, calculating longitudinal and transverse, course maneuvering when air maneuver
The required rudder face degree of bias；
Step 2: the bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver；
Step 3: the cockpit manipulation displacement of aircraft and steering force during calculating air maneuver.
Step 4: different target when according to air maneuver overload repeats step 1- step 3, when obtaining air maneuver
Rod gradient, the bar displacement gradient that longitudinal and transverse, course cockpit manipulation displacement, steering force overload corresponding target.
The bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver described in step 2, is also wrapped
Include following steps:
Step 2.1: determining aircraft along fuselage linear deformation and aircraft along wing transversely deforming according to flight status；
Step 2.2: determine each installation point of aircraft mechanical manoeuvring system, shaft fulcrum with aircraft along fuselage linear deformation and
Spatial position change amount of the aircraft after wing transversely deforming；
Step 2.3: calculating machine steerable system two adjacent two phases between installation point, shaft fulcrum aboard
To deformation.
Step 2.4: by mechanical manoeuvring system aboard all adjacent two it is opposite between installation point, shaft fulcrum
Deformation summation, it is final to determine that steerable system in the total deformation of current flight state, that is, determines and deviates steerable system design theory value
Bias；
Step 2.5: the mechanical manoeuvring system number under current maneuver state is corrected in the deviation being calculated according to step 2.4
According to.
Cockpit manipulation displacement and the steering force that aircraft is calculated described in step 3, specially fly according to what is obtained in step 1
The rudder face degree of bias needed for longitudinal and transverse, course maneuvering when machine is motor-driven and step 2 calculate determine current maneuver during under different overloads
The bias that mechanical manoeuvring system is deformed relative to aircraft is corrected aircraft mechanical manoeuvring system true transmission ratio in real time, is calculated
The cockpit manipulation displacement of aircraft and steering force, and calculate aircraft portrait rod gradient, bar displacement gradient.
According to the requirement of resultant force and resultant moment when air maneuver, rudder needed for longitudinal and transverse, course maneuvering when calculating air maneuver
The face degree of bias；Main state of flight in full flight profile, mission profile determined by overall aircraft master-plan, the weight including aircraft are needed at this time
Center of gravity condition, flap state, engine condition, flying speed, flying height are measured, according to the pneumatic number for considering aircraft flexible deformation
According to, Aircraft Quality characteristic, dynamic characteristics data, then longitudinal and transverse course when calculating aircraft balanced balance needed for the rudder face degree of bias.
The mechanical handling system under current maneuver state is corrected in the deviation being calculated described in step 2.5 according to step 2.4
System data, the mechanical manoeuvring system data are revised true transmission ratio.
Beneficial effects of the present invention: provide a kind of aircraft and mechanical manoeuvring system uneven deformation influences to design on manoeuvre
Method prevention and prevents aircraft so as to ensure that the full mission profile of aircraft meets air maneuver handling characteristic design requirement
Occur non-instruction sexual deviation when friction speed is motor-driven, realizes aircraft true aircraft mechanical manoeuvring system characteristic awing
Determine that process, and true air maneuver handling characteristic design the method for determination, thus ensure that the consistency of flight and design,
Mechanical manoeuvring system characteristic error is had modified, while ensure that for example big table speed of part flight envelope boundary point, big overload flight shape
The precision and accuracy of the manoeuvre design of state, improve the safety of aircraft flight, improve handling quality, meanwhile, this hair
Bright also to have system full and accurate, efficiency is higher；It gives prominence to the key points, considers that comprehensively, accuracy is high；The characteristics of strong applicability.
Detailed description of the invention
Fig. 1 is vector mechanical manoeuvring system schematic diagram；
Fig. 2 is that aircraft positive g flight course middle fuselage upper and lower surface deforms schematic diagram.
Specific embodiment
Aircraft and mechanical manoeuvring system uneven deformation influence design method to manoeuvre, comprising the following steps:
Step 1: according to the requirement of resultant force and resultant moment when air maneuver, calculating longitudinal and transverse, course maneuvering when air maneuver
The required rudder face degree of bias；
Step 2: the bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver；
Step 3: the cockpit manipulation displacement of aircraft and steering force during calculating air maneuver.
Step 4: different target when according to air maneuver overload repeats step 1- step 3, when obtaining air maneuver
Rod gradient, the bar displacement gradient that longitudinal and transverse, course cockpit manipulation displacement, steering force overload corresponding target.
The bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver described in step 2, is also wrapped
Include following steps:
Step 2.1: determining aircraft along fuselage linear deformation and aircraft along wing transversely deforming according to flight status；
Step 2.2: determine each installation point of aircraft mechanical manoeuvring system, shaft fulcrum with aircraft along fuselage linear deformation and
Spatial position change amount of the aircraft after wing transversely deforming；
Step 2.3: calculating machine steerable system two adjacent two phases between installation point, shaft fulcrum aboard
To deformation.
Step 2.4: by mechanical manoeuvring system aboard all adjacent two it is opposite between installation point, shaft fulcrum
Deformation summation, it is final to determine that steerable system in the total deformation of current flight state, that is, determines and deviates steerable system design theory value
Bias；
Step 2.5: the mechanical manoeuvring system number under current maneuver state is corrected in the deviation being calculated according to step 2.4
According to.
Cockpit manipulation displacement and the steering force that aircraft is calculated described in step 3, specially fly according to what is obtained in step 1
The rudder face degree of bias needed for longitudinal and transverse, course maneuvering when machine is motor-driven and step 2 calculate determine current maneuver during under different overloads
The bias that mechanical manoeuvring system is deformed relative to aircraft is corrected aircraft mechanical manoeuvring system true transmission ratio in real time, is calculated
The cockpit manipulation displacement of aircraft and steering force, and calculate aircraft portrait rod gradient, bar displacement gradient.
According to the requirement of resultant force and resultant moment when air maneuver, rudder needed for longitudinal and transverse, course maneuvering when calculating air maneuver
The face degree of bias；Main state of flight in full flight profile, mission profile determined by overall aircraft master-plan, the weight including aircraft are needed at this time
Center of gravity condition, flap state, engine condition, flying speed, flying height are measured, according to the pneumatic number for considering aircraft flexible deformation
According to, Aircraft Quality characteristic, dynamic characteristics data, then longitudinal and transverse course when calculating aircraft balanced balance needed for the rudder face degree of bias.
The mechanical handling system under current maneuver state is corrected in the deviation being calculated described in step 2.5 according to step 2.4
System data, the mechanical manoeuvring system data are revised true transmission ratio.
Embodiment:
Scheme of installation is shown in Fig. 1 to vector mechanical manoeuvring system aboard, and aircraft mechanical manoeuvring system is aboard
Whole deformation amount should include aircraft body and mechanical manoeuvring system relative deformation between each installation point, shaft fulcrum
Combination superposition.
Flexible deformation can occur in flight for airframe, and motor-driven middle aircraft deflection can be with mistake in air maneuver
The variation of load and change.Motor-driven for aircraft portrait, if normal g-load is positive g, fuselage is bent downwardly, this opportunity
The portion of carrying on the body can elongate.Fuselage can generate the flexible deformation shaped like " shoulder pole ", see Fig. 2, and the free end in Fig. 2 is that aircraft is a certain transversal
Face, the fuselage back between two cross sections can elongate generation amount of elongation, and bottom can compress generation shortening amount.Assuming that aircraft is mechanical
The two neighboring installation point of the steerable system free end in Fig. 2 respectively, this section mechanical manoeuvring system leverage relative to machine
Body generates relative deformation, and mechanical manoeuvring system leverage is that multiple mechanical shafts are formed by connecting, and general aircraft is frequently with mechanical handling
System leverage major part is located at fuselage back (dorsal fin position), when fuselage back elongates, since the rigidity of steerable system is remote
Greater than the rigidity of airframe, control linkage will not change, the mechanical handling leverage of fuselage back with the variation of fuselage length
Main rocker shaft fulcrum coordinate position changes, at this point, control linkage is shortened relative to fuselage, this will lead to bar
The Forward of head position is to generate steerable system to " deformation " for aircraft.Similarly, machinery can be also generated during aircraft rolling
The relative deformation of installation site of the steerable system in wing.
In the design, aircraft and mechanical manoeuvring system should all be designed according to elastomer respectively, but on the whole, mechanical
Steerable system rigidity Design is larger, moreover, only the frictional force by cockpit steering force, servomechanism control valve (is manipulated without power is returned
System), the hinge moment (resilient steerable system) of rudder face, system variant is relatively small.Elastic Aircraft and mechanical manoeuvring system
The equation of motion is the rigid motion and elastic vibration freedom degree for comprehensively considering aircraft and mechanical manoeuvring system, from the angle of system capacity
Degree sets out, and the general equation of motion of Elastic Aircraft and mechanical manoeuvring system movement is derived using the Lagrange equation of motion.
If each elastic vibration mode has acquired, and is grasped with the unit vector of reference body shafting and Elastic Aircraft and machinery
The vertical undeformed shape representation elastic vibration Mode Shape of system motion,
With the increase of normal g-load, the main rocker shaft fulcrum coordinate position of the fore-and-aft control leverage of fuselage back is sent out
Changing.It when normal g-load is positive g, can be obtained according to fuselage deflection data △ H, fuselage generates bending angle △ α:
In formula, △ L is fuselage length deflection, and the distance of fuselage upper surface to bending neutral layer is Ru, therefore, this region
Fuselage upper surface amount of elongation are as follows:
Δ L=RuΔα1
In formula, △ α1Bending angle, course behaviour are generated between a certain section of two neighboring installation point of fuselage and shaft fulcrum
The ratio of vertical system input displacement and output displacement is K1, normal g-load be positive g when, due to steerable system be mounted on it is winged
Machine dorsal fin position (at fuselage upper surface), so leverage causes the displacement of head with respect to the shortening of fuselage are as follows:
ΔXr=Δ L/K1
The rudder rudder face degree of bias and the transmission ratio of head displacement are K in directional control system2, in this way, △ Xr head is displaced
The bias of the amount of rudder of generation is △ δr:
Δδr=Δ Xr×K2
It is also possible to according to the cockpit actuated position of the rudder face position retrospectively calculate aircraft of aircraft.Basis flies at this time simultaneously
The motor-driven required control surface deflection position of machine can calculate cockpit manipulation displacement and steering force when air maneuver.
When aircraft big overload flight, should be deformed respectively according to different the get off the plane bodies and mechanical manoeuvring system itself of overloading
Difference calculate the corresponding cockpit manipulated variable of aircraft, including steering force and manipulation displacement, to show that aircraft is motor-driven in big overload
When true rod gradient and bar displacement gradient.
Claims (5)
1. aircraft and mechanical manoeuvring system uneven deformation influence design method to manoeuvre, it is characterised in that: including following step
It is rapid:
Step 1: according to the requirement of resultant force and resultant moment when air maneuver, needed for longitudinal and transverse, course maneuvering when calculating air maneuver
The rudder face degree of bias；
Step 2: the bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver；
Step 3: the cockpit manipulation displacement of aircraft and steering force during calculating air maneuver；
Step 4: different target when according to air maneuver overload repeats step 1- step 3, when obtaining air maneuver it is vertical,
Rod gradient, the bar displacement gradient that horizontal, course cockpit manipulation displacement, steering force overload corresponding target.
2. aircraft according to claim 1 and mechanical manoeuvring system uneven deformation influence design method to manoeuvre,
Be characterized in that: the bias that mechanical manoeuvring system is deformed relative to aircraft during calculating air maneuver described in step 2 is also wrapped
Include following steps:
Step 2.1: determining aircraft along fuselage linear deformation and aircraft along wing transversely deforming according to flight status；
Step 2.2: determining each installation point of aircraft mechanical manoeuvring system, shaft fulcrum with aircraft along fuselage linear deformation and aircraft
Spatial position change amount after wing transversely deforming；
Step 2.3: the two adjacent two opposite changes between installation point, shaft fulcrum aboard of calculating machine steerable system
Shape；
Step 2.4: by mechanical manoeuvring system all two adjacent relative deformations between installation point, shaft fulcrum aboard
Summation, it is final to determine that steerable system in the total deformation of current flight state, that is, determines and deviates the inclined of steerable system design theory value
From amount；
Step 2.5: the mechanical manoeuvring system data under current maneuver state are corrected in the deviation being calculated according to step 2.4.
3. aircraft according to claim 1 and mechanical manoeuvring system uneven deformation influence design method to manoeuvre,
It is characterized in that: calculating cockpit manipulation displacement and the steering force of aircraft described in step 3, specially according to being obtained in step 1
The rudder face degree of bias needed for longitudinal and transverse, course maneuvering when air maneuver and step 2 calculate determine current maneuver during different overloads
Aircraft mechanical manoeuvring system true transmission ratio in real time, meter are corrected in the bias that lower mechanical manoeuvring system is deformed relative to aircraft
Cockpit manipulation displacement and the steering force of aircraft are calculated, and calculates aircraft portrait rod gradient, bar displacement gradient.
4. aircraft according to claim 1 and mechanical manoeuvring system uneven deformation influence design method to manoeuvre,
It is characterized in that: according to the requirement of resultant force and resultant moment when air maneuver, needed for longitudinal and transverse, course maneuvering when calculating air maneuver
The rudder face degree of bias；Main state of flight in full flight profile, mission profile determined by overall aircraft master-plan is needed at this time, including aircraft
Weight center of gravity state, flap state, engine condition, flying speed, flying height, according to the pneumatic of consideration aircraft flexible deformation
Data, Aircraft Quality characteristic, dynamic characteristics data, then longitudinal and transverse course when calculating aircraft balanced balance needed for the rudder face degree of bias.
5. aircraft according to claim 2 and mechanical manoeuvring system uneven deformation influence design method to manoeuvre,
Be characterized in that: the mechanical handling under current maneuver state is corrected in the deviation being calculated described in step 2.5 according to step 2.4
System data, the mechanical manoeuvring system data are revised true transmission ratio.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110920933A (en) * | 2019-12-04 | 2020-03-27 | 中国直升机设计研究所 | Helicopter control lever frequency modulation design method |
| CN117390774A (en) * | 2023-12-13 | 2024-01-12 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
Citations (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1480652A (en) * | 1974-01-02 | 1977-07-20 | Gen Electric | Heading reference system for an aircraft |
| ES8204681A1 (en) * | 1980-04-01 | 1982-05-16 | United Technologies Corp | Multi-axis force stick, self- trimmed aircraft flight control system |
| CN1175017A (en) * | 1996-08-22 | 1998-03-04 | 波音公司 | Aircraft pitch-axis stability and command augmentation system |
| US5738310A (en) * | 1994-12-22 | 1998-04-14 | Eurocopter France | Rudder bar system with force gradient for a helicopter |
| US20080255713A1 (en) * | 2007-04-13 | 2008-10-16 | Dan Onu | Dynamic Adjustment of Wing Surfaces for Variable Camber |
| JP2011230555A (en) * | 2010-04-23 | 2011-11-17 | Jtekt Corp | Steering device for aircraft |
| US20120072056A1 (en) * | 2008-03-31 | 2012-03-22 | Muhammad Emadul Hasan | Flight control system for rotary wing aircraft |
| US20150028162A1 (en) * | 2011-06-10 | 2015-01-29 | Eads Deutschland Gmbh | Method and apparatus for minimizing dynamic structural loads of an aircraft |
| CN105383684A (en) * | 2015-12-12 | 2016-03-09 | 中国航空工业集团公司西安飞机设计研究所 | Compensation control method for asymmetric thrust of plane |
| CN105819001A (en) * | 2016-05-12 | 2016-08-03 | 中国航空工业集团公司西安飞机设计研究所 | Determination method for horizontal tail installation angle of horizontal tail fixed airplane |
| CN105892482A (en) * | 2016-04-01 | 2016-08-24 | 哈尔滨飞机工业集团有限责任公司 | Trimming adapter of automatic flight control system |
| CN106020211A (en) * | 2016-05-12 | 2016-10-12 | 中国航空工业集团公司西安飞机设计研究所 | Method for calculating preset skewness of elevator at taking-off time of airplane |
| CN106091964A (en) * | 2016-05-31 | 2016-11-09 | 中国航空工业集团公司西安飞机设计研究所 | A kind of pull bar interlinked mechanism connector strain measurement method |
| CN106184715A (en) * | 2015-04-29 | 2016-12-07 | 陕西飞机工业（集团）有限公司 | A kind of aircraft control stick trimming system |
| CN106530897A (en) * | 2016-12-28 | 2017-03-22 | 中国航空工业集团公司西安飞机设计研究所 | Flight simulation training device |
| CN106649909A (en) * | 2016-08-29 | 2017-05-10 | 成都飞机工业（集团）有限责任公司 | Dual-redundancy compensation type empennage control plane fault state control method |
| CN106681350A (en) * | 2017-02-22 | 2017-05-17 | 北京经纬恒润科技有限公司 | Trimming method and device for aircraft system |
| US20170361941A1 (en) * | 2016-06-16 | 2017-12-21 | Ge Aviation Systems Llc | Controlling aircraft using thrust differential trim |
-
2018
- 2018-11-02 CN CN201811305733.0A patent/CN109592064B/en active Active
Patent Citations (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1480652A (en) * | 1974-01-02 | 1977-07-20 | Gen Electric | Heading reference system for an aircraft |
| ES8204681A1 (en) * | 1980-04-01 | 1982-05-16 | United Technologies Corp | Multi-axis force stick, self- trimmed aircraft flight control system |
| US5738310A (en) * | 1994-12-22 | 1998-04-14 | Eurocopter France | Rudder bar system with force gradient for a helicopter |
| CN1175017A (en) * | 1996-08-22 | 1998-03-04 | 波音公司 | Aircraft pitch-axis stability and command augmentation system |
| US20080255713A1 (en) * | 2007-04-13 | 2008-10-16 | Dan Onu | Dynamic Adjustment of Wing Surfaces for Variable Camber |
| US20120072056A1 (en) * | 2008-03-31 | 2012-03-22 | Muhammad Emadul Hasan | Flight control system for rotary wing aircraft |
| JP2011230555A (en) * | 2010-04-23 | 2011-11-17 | Jtekt Corp | Steering device for aircraft |
| US20150028162A1 (en) * | 2011-06-10 | 2015-01-29 | Eads Deutschland Gmbh | Method and apparatus for minimizing dynamic structural loads of an aircraft |
| CN106184715A (en) * | 2015-04-29 | 2016-12-07 | 陕西飞机工业（集团）有限公司 | A kind of aircraft control stick trimming system |
| CN105383684A (en) * | 2015-12-12 | 2016-03-09 | 中国航空工业集团公司西安飞机设计研究所 | Compensation control method for asymmetric thrust of plane |
| CN105892482A (en) * | 2016-04-01 | 2016-08-24 | 哈尔滨飞机工业集团有限责任公司 | Trimming adapter of automatic flight control system |
| CN106020211A (en) * | 2016-05-12 | 2016-10-12 | 中国航空工业集团公司西安飞机设计研究所 | Method for calculating preset skewness of elevator at taking-off time of airplane |
| CN105819001A (en) * | 2016-05-12 | 2016-08-03 | 中国航空工业集团公司西安飞机设计研究所 | Determination method for horizontal tail installation angle of horizontal tail fixed airplane |
| CN106091964A (en) * | 2016-05-31 | 2016-11-09 | 中国航空工业集团公司西安飞机设计研究所 | A kind of pull bar interlinked mechanism connector strain measurement method |
| US20170361941A1 (en) * | 2016-06-16 | 2017-12-21 | Ge Aviation Systems Llc | Controlling aircraft using thrust differential trim |
| CN106649909A (en) * | 2016-08-29 | 2017-05-10 | 成都飞机工业（集团）有限责任公司 | Dual-redundancy compensation type empennage control plane fault state control method |
| CN106530897A (en) * | 2016-12-28 | 2017-03-22 | 中国航空工业集团公司西安飞机设计研究所 | Flight simulation training device |
| CN106681350A (en) * | 2017-02-22 | 2017-05-17 | 北京经纬恒润科技有限公司 | Trimming method and device for aircraft system |
Non-Patent Citations (7)
| Title |
|---|
| 卢京明: "飞机全机主操纵系统疲劳试验故障诊断与分析", 《飞机设计》 * |
| 吕新波等: "民用飞机俯仰操纵力特性设计研究", 《飞行力学》 * |
| 吕玉虎等: "飞机纵向振荡的试飞研究", 《飞行力学》 * |
| 屈宗源: "某型飞机升降舵操纵系统不平衡力问题的分析与解决", 《科技创新》 * |
| 席锋等: "机械操纵系统飞机航向偏离现象分析", 《测控技术》 * |
| 梁东山等: "BW-1变稳飞机地面试验研究", 《飞行力学》 * |
| 王旻: "民机电传飞控系统适航要求与符合性方法研究", 《中国民航大学学报》 * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110920933A (en) * | 2019-12-04 | 2020-03-27 | 中国直升机设计研究所 | Helicopter control lever frequency modulation design method |
| CN117390774A (en) * | 2023-12-13 | 2024-01-12 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
| CN117390774B (en) * | 2023-12-13 | 2024-03-19 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
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