CN106528915B - Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane - Google Patents

Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane Download PDF

Info

Publication number
CN106528915B
CN106528915B CN201610837895.3A CN201610837895A CN106528915B CN 106528915 B CN106528915 B CN 106528915B CN 201610837895 A CN201610837895 A CN 201610837895A CN 106528915 B CN106528915 B CN 106528915B
Authority
CN
China
Prior art keywords
rate
inf
lim
max
control surface
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201610837895.3A
Other languages
Chinese (zh)
Other versions
CN106528915A (en
Inventor
武虎子
吕新波
黄振威
伍智敏
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Xian Aircraft Design and Research Institute of AVIC
Original Assignee
Xian Aircraft Design and Research Institute of AVIC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Xian Aircraft Design and Research Institute of AVIC filed Critical Xian Aircraft Design and Research Institute of AVIC
Priority to CN201610837895.3A priority Critical patent/CN106528915B/en
Publication of CN106528915A publication Critical patent/CN106528915A/en
Application granted granted Critical
Publication of CN106528915B publication Critical patent/CN106528915B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • Geometry (AREA)
  • General Physics & Mathematics (AREA)
  • Evolutionary Computation (AREA)
  • General Engineering & Computer Science (AREA)
  • Computer Hardware Design (AREA)
  • Automation & Control Theory (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Computational Mathematics (AREA)
  • Mathematical Analysis (AREA)
  • Mathematical Optimization (AREA)
  • Pure & Applied Mathematics (AREA)
  • Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)

Abstract

The invention discloses a method for determining a minimum deflection rate critical value of a maneuvering control surface of a civil transport plane. The method comprises the following steps: step 1: calculating the roll rate p at any moment in the flying state of the airplanei_infPitch angle rate qi_infYaw rate ri_inf(ii) a Step 2: respectively calculating the roll angle rate, the pitch angle rate and the yaw angle rate at any moment; and step 3: calculating KMaximum euler angular rate when inf and calculate KMaximum euler angular rate when Lim; and 4, step 4: respectively recording the time t required for the roll rate, the pitch angle rate and the yaw rate to reach the maximum values1And tjAnd pass t1And tjDetermining the time difference dtj(ii) a And 5: a minimum rate of deflection threshold for the motorized control surface is determined. The method for determining the minimum deflection rate critical value of the maneuvering control surface of the civil transport engine provides a set of complete method for determining the minimum deflection rate critical value of the maneuvering control surface of the civil transport engine; provides a new means for designing the deflection rate of the maneuvering surface.

Description

Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane
Technical Field
The invention relates to the technical field of aircraft stability design, in particular to a method for determining a minimum deflection rate critical value of a maneuvering control surface of a civil transport plane.
Background
The minimum deflection rate critical value of the maneuvering control surface of the airplane can ensure the maneuvering capability and the flight safety of the airplane, if the deflection rate is too small, the expected maneuvering capability of the airplane can not be ensured, and even the flight safety is threatened; the minimum yaw rate threshold determination is particularly important because excessive actuator design requirements may be imposed at a cost if the yaw rate is excessive. FIG. 1 is a schematic diagram of an aircraft maneuver control surface minimum yaw rate threshold determination.
In FIG. 1, the solid line represents KEuler angular rate curve at inf;the dotted line represents KEuler angular rate curve at Lim; t is t1Is represented by KTime required to reach maximum euler angular rate at inf; t is tjIs represented by KTime required to reach maximum euler angular rate when Lim; dtjRepresenting the time difference (t)j-t1)。
At present, no suitable method exists for determining the minimum deflection rate critical value of the maneuvering surface of the civil transport plane, most of the methods are determined by referring to the minimum deflection rate of the maneuvering surface of the similar plane at home and abroad, but the minimum deflection rate is determined so and is large, theoretical technical support is lacked, and even the minimum deflection rate requirement is met so that the development cost of the plane is greatly increased.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
It is an object of the present invention to provide a method for determining a threshold minimum yaw rate for a mobile control surface of a civil transport vehicle that overcomes or at least mitigates at least one of the above-mentioned disadvantages of the prior art.
In order to achieve the above object, the present invention provides a method for determining a minimum deflection rate threshold value of a civil transporter maneuvering control surface, comprising the following steps:
step 1: on a motorized control surface KWhen the value is inf, the roll rate p of the airplane at any moment in the flying state is calculatedi_infPitch angle rate qi_infYaw rate ri_inf；
Step 2: when the maneuvering surface is KRespectively calculating the roll rate (p) at any moment when the series of finite values Lim are obtainedi_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)j；
And step 3: calculating KMaximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) And calculating KMaximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j)；
And 4, step 4: respectively recording K by means of computational simulationInf and Lim, the time t required for the roll rate, pitch rate, and yaw rate to reach maximum values1And tjAnd pass t1And tjDetermining the time difference dtj；
And 5: determining a minimum rate of deflection threshold (K) for a motorized control surface_min)L。
Preferably, the roll rate p in step 1i_infPitch angle rate qi_infYaw rate ri_infSpecifically, the method is obtained by the following formula:
Figure GDA0002546319700000021
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia about the three axes x, y, z of the aircraft, kg.m3，pi-1_inf,qi-1_inf,ri-1_infIs the euler angular rate, deg/s,
Figure GDA0002546319700000022
is the Euler angular acceleration at the previous moment, ° s2Δ t is the calculation time step, s, takes the value of 0.001,
Figure GDA0002546319700000023
is the rapid pressure at the previous moment in kg/m.s2，
Figure GDA0002546319700000024
The roll, pitch and yaw moment coefficients at the previous moment are respectively, S is the wing reference area, m2，bA,cAIs the wing span length and the average aerodynamic chord length, m.
Preferably, the roll rate (p) in step 2i_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)jSpecifically, the method is obtained by the following formula:
Figure GDA0002546319700000031
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia about the three axes x, y, z of the aircraft, kg.m3，((pi-1_Lim)j,(qi-1_Lim)j,(ri-1_Lim)j) Is the euler angular rate, deg/s,
Figure GDA0002546319700000032
is the Euler angular acceleration at the previous moment, ° s2Δ t is the calculation time step, s, takes the value of 0.001,
Figure GDA0002546319700000033
is the rapid pressure at the previous moment in kg/m.s2，
Figure GDA0002546319700000034
The roll, pitch and yaw moment coefficients at the previous moment are respectively, S is the wing reference area, m2，bA,cAIs the wing span length and the average aerodynamic chord length, m.
Preferably, K in said step 3Maximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) The calculation is made by the following formula:
pmax_inf＝max(pi_inf),qmax_inf＝max(qi_inf),rmax_inf＝max(ri_inf) (ii) a Wherein the content of the first and second substances,
i is an arbitrary time.
Preferably, K in said step 3Maximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j) The calculation is made by the following formula:
(pmax_Lim)j＝max((pi_Lim)j),(qmax_Lim)j＝max((qi_Lim)j),(rmax_Lim)j＝max((ri_Lim)j) (ii) a Wherein the content of the first and second substances,
j is KThe number of values taken.
Preferably, said time difference dtjIs dtj＝tj-t1。
Preferably, the step 5 is: judgment of dtjSeveral time difference values satisfy the following conditions: the first preset condition is not more than dtjNot more than a second preset condition;
if only one time difference value meets the condition, calculating and simulating to record the current control surface deflection value through the time difference value, wherein the current control surface deflection value is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L；
If a plurality of time difference values meet the condition, the time difference value with the numerical value closest to the first preset condition in the plurality of time difference values is taken, the current control surface deflection value is calculated and recorded in a simulation mode according to the time difference value, and the current control surface deflection value is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L。
Preferably, the first preset condition is 0.075 seconds; the second preset condition is 0.08 seconds.
Preferably, the series of finite values (Lim) in step 2 includes 5 °/s,10 °/s,15 °/s,20 °/s,25 °/s, 30 °/s.
The method for determining the minimum deflection rate critical value of the maneuvering control surface of the civil transport engine provides a set of complete method for determining the minimum deflection rate critical value of the maneuvering control surface of the civil transport engine; provides a new means for the design of the deflection rate of the maneuvering surface, and simultaneously better serves the design of future models.
Drawings
FIG. 1 is a schematic diagram of a prior art determination of a minimum yaw rate threshold for an aircraft maneuver surface.
Fig. 2 is a flow chart illustrating a method for determining a threshold value of minimum yaw rate of a maneuvering surface of a civil transport vehicle according to an embodiment of the invention.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present invention clearer, the technical solutions in the embodiments of the present invention will be described in more detail below with reference to the accompanying drawings in the embodiments of the present invention. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are only some, but not all embodiments of the invention. The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention. Embodiments of the present invention will be described in detail below with reference to the accompanying drawings.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present invention and for simplifying the description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the scope of the present invention.
FIG. 1 is a schematic diagram of a prior art determination of a minimum yaw rate threshold for an aircraft maneuver surface. Fig. 2 is a flow chart illustrating a method for determining a threshold value of minimum yaw rate of a maneuvering surface of a civil transport vehicle according to an embodiment of the invention.
The method for determining the minimum deflection rate threshold value of the maneuvering surface of a civil transport aircraft as shown in fig. 1 comprises the following steps: step 1: on a motorized control surface KWhen the value is inf, the roll rate p of the airplane at any moment in the flying state is calculatedi_infPitch angle rate qi_infYaw rate ri_inf；
Step 2: when the maneuvering surface is KRespectively calculating the roll rate (p) at any moment when the series of finite values Lim are obtainedi_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)j；
And step 3: calculating KMaximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) And calculating KMaximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j)；
And 4, step 4: respectively recording K by means of computational simulationInf and Lim, the time t required for the roll rate, pitch rate, and yaw rate to reach maximum values1And tjAnd pass t1And tjDetermining the time difference dtj；
And 5: determining a minimum rate of deflection threshold (K) for a motorized control surface_min)L。
In the present embodiment, the roll rate p in step 1i_infPitch angle rate qi_infYaw rate ri_infSpecifically, the method is obtained by the following formula:
Figure GDA0002546319700000051
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia (in kg. m) about the three axes x, y, z of the aircraft3)(xIn correspondence with the x-axis,yin correspondence with the y-axis,zcorresponding to the z-axis), pi-1_inf,qi-1_inf,ri-1_infIs the euler angular rate (in deg/s) at the previous time,
Figure GDA0002546319700000061
respectively, the euler angular acceleration (in deg/s) of the last moment2) And delta t is the calculation time step (unit is s) and takes the value of 0.001,
Figure GDA0002546319700000062
is the rapid pressure (in kg/m.s.) at the previous moment2)，
Figure GDA0002546319700000063
Is the roll coefficient at the last moment,
Figure GDA0002546319700000064
Is the pitch coefficient at the previous moment,
Figure GDA0002546319700000065
Is the yaw moment coefficient at the previous moment, S is the wing reference area (in m)2)，bAIs the span length (m), c of the wingAMean aerodynamic chord length (in m).
In this embodiment, the roll rate (p) in step 2i_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)jSpecifically, the method is obtained by the following formula:
Figure GDA0002546319700000066
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia (in kg. m) about the three axes x, y, z of the aircraft3)(xIn correspondence with the x-axis,yin correspondence with the y-axis,zcorresponding to the z-axis) ((p)i-1_Lim)j，(qi-1_Lim)j，(ri-1_Lim)j) Is the euler angular rate (in deg/s) at the previous time,
Figure GDA0002546319700000067
respectively, the euler angular acceleration (in deg/s) of the last moment2) And delta t is the calculation time step (unit is s) and takes the value of 0.001,
Figure GDA0002546319700000068
is the rapid pressure (in kg/m.s.) at the previous moment2)，
Figure GDA0002546319700000069
Is the roll coefficient at the last moment,
Figure GDA00025463197000000611
Is the pitch coefficient at the previous moment,
Figure GDA00025463197000000610
Is the yaw moment coefficient at the previous moment, S is the wing reference area (in m)2)，bAIs the span length (m), c of the wingAMean aerodynamic chord length (in m).
In this embodiment, K in step 3Maximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) The calculation is made by the following formula:
pmax_inf＝max(pi_inf),qmax_inf＝max(qi_inf),rmax_inf＝max(ri_inf) (ii) a Wherein the content of the first and second substances,
i is an arbitrary time.
In this embodiment, K in step 3Maximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j) The calculation is made by the following formula:
(pmax_Lim)j＝max((pi_Lim)j),(qmax_Lim)j＝max((qi_Lim)j),(rmax_Lim)j＝max((ri_Lim)j) (ii) a Wherein j is KThe number of values taken.
In the present embodiment, the time difference dtjIs dtj＝tj-t1。
In this embodiment, step 5 is: judgment of dtjSeveral time difference values satisfy the following conditions:
the first preset condition is not more than dtjNot more than a second preset condition;
if only one time difference satisfies the condition, the time is passedCalculating and simulating to record the current control surface deflection value which is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L；
If a plurality of time difference values meet the condition, the time difference value with the numerical value closest to the first preset condition in the plurality of time difference values is taken, the current control surface deflection value is calculated and recorded in a simulation mode according to the time difference value, and the current control surface deflection value is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L。
In the present embodiment, the first preset condition is 0.075 seconds; the second preset condition is 0.08 seconds.
In this embodiment, the series of limited values Lim in step 2 includes 5 °/s,10 °/s,15 °/s,20 °/s,25 °/s, and 30 °/s. It is understood that a series of limited Lim values may be selected at will.
The present application is described in further detail below by way of examples, it being understood that the examples do not constitute any limitation to the present application.
Given the aircraft calculation state of one embodiment, table 1 below:
Figure GDA0002546319700000071
by the method for determining the minimum deflection rate critical value of the maneuvering control surface of the civil transport aircraft, the following calculation results can be obtained:
TABLE 2 minimum yaw rate thresholds for motorized control surfaces (elevators)
Figure GDA0002546319700000081
TABLE 3 minimum yaw rate thresholds for motorized control surfaces (rudders)
Figure GDA0002546319700000082
TABLE 4 minimum deflection Rate thresholds for motorized control surfaces (ailerons)
Figure GDA0002546319700000083
Figure GDA0002546319700000091
Finally, it should be pointed out that: the above examples are only for illustrating the technical solutions of the present invention, and are not limited thereto. Although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.

Claims (7)

1. A method for determining a minimum deflection rate threshold value of a civil transport vehicle maneuvering control surface is characterized by comprising the following steps:
step 1: on a motorized control surface KWhen the value is inf, the roll angle rate p at any moment in the flight state of the airplane is calculated through a formulai_infPitch angle rate qi_infYaw rate ri_inf；
Step 2: when the maneuvering surface is KFor a series of limited values Lim, the roll rate (p) at any time is calculated by a formulai_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)j；
And step 3: calculating KMaximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) And calculating KMaximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j)；
And 4, step 4: respectively recording K by means of computational simulationInf and Lim, the time t required for the roll rate, pitch rate, and yaw rate to reach maximum values1And tjAnd pass t1And tjDetermining the time difference dtjSaid time difference dtjIs dtj＝tj-t1；
And 5: determining a minimum rate of deflection threshold (K) for a motorized control surface_min)LThe method specifically comprises the following steps:
judgment of dtjSeveral time difference values satisfy the following conditions: the first preset condition is not more than dtjNot more than a second preset condition;
if only one time difference value meets the condition, calculating and simulating to record the current control surface deflection value through the time difference value, wherein the current control surface deflection value is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L；
If a plurality of time difference values meet the condition, the time difference value with the numerical value closest to the first preset condition in the plurality of time difference values is taken, the current control surface deflection value is calculated and recorded in a simulation mode according to the time difference value, and the current control surface deflection value is the minimum deflection rate critical value (K) of the maneuvering control surface_min)L。
2. The method for determining threshold value of minimum yaw rate of mobile control surface of civil transport vehicle as defined in claim 1, wherein the roll rate p in step 1i_infPitch angle rate qi_infYaw rate ri_infSpecifically, the method is obtained by the following formula:
Figure FDA0002546319690000021
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia about the three axes x, y, z of the aircraft, kg.m3，pi-1_inf,qi-1_inf,ri-1_infIs the euler angular rate, deg/s,
Figure FDA0002546319690000022
is the Euler angular acceleration at the previous moment, ° s2Δ t is the calculation time step, s, takes the value of 0.001,
Figure FDA0002546319690000023
is the rapid pressure at the previous moment in kg/m.s2，
Figure FDA0002546319690000024
The roll, pitch and yaw moment coefficients at the previous moment are respectively, S is the wing reference area, m2，bA,cAIs the wing span length and the average aerodynamic chord length, m.
3. Method for determining the threshold value of the minimum yaw rate of a mobile control surface of a civil transport vehicle according to claim 2, characterised in that the roll rate (p) in step 2i_Lim)jPitch angle rate (q)i_Lim)jYaw rate (r)i_Lim)jSpecifically, the method is obtained by the following formula:
Figure FDA0002546319690000025
wherein the content of the first and second substances,
Ix,Iy,Izis the moment of inertia about the three axes x, y, z of the aircraft, kg.m3,((pi-1_Lim)j，(qi-1_Lim)j，(ri-1_Lim)j) Is the euler angular rate, deg/s,
Figure FDA0002546319690000026
respectively euler angular acceleration at the previous moment, degree/s 2, delta t is a calculation time step length, s is 0.001,
Figure FDA0002546319690000031
is the rapid pressure at the previous moment in kg/m.s2，
Figure FDA0002546319690000032
The roll, pitch and yaw moment coefficients at the previous moment are respectively, S is the wing reference area, m2，bA,cAIs the wing span length and the average aerodynamic chord length, m.
4. Method for determining the threshold value of minimum deflection rate of a mobile control surface of a civil transport vehicle according to claim 3, characterised in that K in step 3Maximum euler angular rate (p) at infmax_inf，qmax_inf，rmax_inf) The calculation is made by the following formula:
pmax_inf＝max(pi_inf),qmax_inf＝max(qi_inf),rmax_inf＝max(ri_inf) (ii) a Wherein the content of the first and second substances,
i is an arbitrary time.
5. Method for determining the threshold value of minimum deflection rate of a mobile control surface of a civil transport vehicle according to claim 4, characterised in that K in step 3Maximum euler angular rate when Lim ((p)max_Lim)j，(qmax_Lim)j，(rmax_Lim)j) The calculation is made by the following formula:
(pmax_Lim)j＝max((pi_Lim)j),(qmax_Lim)j＝max((qi_Lim)j),(rmax_Lim)j＝max((ri_Lim)j) (ii) a Wherein j is KThe number of values taken.
6. The method for determining a threshold value of minimum yaw rate of a civil transporter maneuvering surface of claim 1, characterized in that the first preset condition is 0.075 seconds; the second preset condition is 0.08 seconds.
7. The method of claim 1, wherein the finite Lim values in step 2 include 5 °/s,10 °/s,15 °/s,20 °/s,25 °/s, and 30 °/s.
CN201610837895.3A 2016-09-21 2016-09-21 Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane Active CN106528915B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201610837895.3A CN106528915B (en) 2016-09-21 2016-09-21 Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201610837895.3A CN106528915B (en) 2016-09-21 2016-09-21 Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane

Publications (2)

Publication Number Publication Date
CN106528915A CN106528915A (en) 2017-03-22
CN106528915B true CN106528915B (en) 2020-08-11

Family

ID=58343933

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201610837895.3A Active CN106528915B (en) 2016-09-21 2016-09-21 Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane

Country Status (1)

Country Link
CN (1) CN106528915B (en)

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110598345B (en) * 2019-09-19 2022-12-16 中国航空工业集团公司西安飞机设计研究所 Method for calculating and analyzing rolling characteristics of fuel in oil tank of oiling machine during transverse movement

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4267731A (en) * 1980-02-11 1981-05-19 Sperry Corporation Force balanced vibratory rate sensor
CN101521558A (en) * 2008-03-01 2009-09-02 华为技术有限公司 Self-adaptive multi-speed speed and phase adjusting method and device therefor
CN101625571A (en) * 2009-07-25 2010-01-13 大连理工大学 Method for simulating six degrees of freedom movement of spinning aircraft
CN102607561A (en) * 2012-02-28 2012-07-25 西安费斯达自动化工程有限公司 Aircraft Euler angle correction model based on accelerometer

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4267731A (en) * 1980-02-11 1981-05-19 Sperry Corporation Force balanced vibratory rate sensor
CN101521558A (en) * 2008-03-01 2009-09-02 华为技术有限公司 Self-adaptive multi-speed speed and phase adjusting method and device therefor
CN101625571A (en) * 2009-07-25 2010-01-13 大连理工大学 Method for simulating six degrees of freedom movement of spinning aircraft
CN102607561A (en) * 2012-02-28 2012-07-25 西安费斯达自动化工程有限公司 Aircraft Euler angle correction model based on accelerometer

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
飞翼布局飞机舵面偏转速率设计;何倩琳等;《北京航空航天大学学报》;20150930;第41卷(第9期);第1729-1735页 *

Also Published As

Publication number Publication date
CN106528915A (en) 2017-03-22

Similar Documents

Publication Publication Date Title
CN106843245B (en) Unmanned aerial vehicle attitude control method and device and unmanned aerial vehicle
CN107065539B (en) Control surface fault self-adaptive fault-tolerant control method of flying wing layout aircraft
CN110502027B (en) Four-rotor unmanned aerial vehicle attitude fault-tolerant control method based on self-adaptive terminal sliding mode
CA2789269C (en) Stall prevention/recovery system and method
US11651699B2 (en) Predictive aircraft flight envelope protection system
CN106649909B (en) Dual-redundancy compensation type empennage control surface fault state control method
He et al. Acceleration-feedback-enhanced robust control of an unmanned helicopter
CN111538255B (en) Anti-bee colony unmanned aerial vehicle aircraft control method and system
CN111367182A (en) Hypersonic aircraft anti-interference backstepping control method considering input limitation
CN105819001B (en) A kind of horizontal tail setting angle of horizontal tail fastening aircraft determines method
CN111443721A (en) Attitude dynamic surface control method for quad-rotor unmanned aerial vehicle and storage medium
CN113485400A (en) Roll control method for vertical launch unmanned aerial vehicle
CN106528915B (en) Method for determining minimum deflection rate critical value of maneuvering control surface of civil transport plane
CN110825115A (en) Limit limit control method for aircraft attack angle and overload
CN105947184A (en) Stability calculation method of incidence angle controller
CN109634110B (en) Tailstock type unmanned aerial vehicle transition section robust control method based on nonlinear disturbance observer
JP2017077882A (en) Roll attitude-dependent roll rate limit
US10099775B2 (en) Autopilot system for an aircraft and related process
CN113156985A (en) Obstacle avoidance robust disturbance rejection flight control method of fixed-wing unmanned aerial vehicle based on preset performance
CN116088549B (en) Tailstock type vertical take-off and landing unmanned aerial vehicle attitude control method
CN109101035B (en) Method for controlling vertical plane trajectory of UUV in high-altitude gliding
CN207902734U (en) A kind of unmanned plane of aerodynamic arrangement
CN106249748B (en) Equipment form control method and controlled motion equipment
CN113885581B (en) Coordinated flight control method and device, electronic equipment and readable storage medium
JP2021536393A (en) Usable time before starting automatic recovery of the aircraft

Legal Events

Date Code Title Description
C06 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant