CN105905296B - Single-blade convolution aircraft lift Optimization Design based on spinning stability - Google Patents
Single-blade convolution aircraft lift Optimization Design based on spinning stability Download PDFInfo
- Publication number
- CN105905296B CN105905296B CN201610329445.3A CN201610329445A CN105905296B CN 105905296 B CN105905296 B CN 105905296B CN 201610329445 A CN201610329445 A CN 201610329445A CN 105905296 B CN105905296 B CN 105905296B
- Authority
- CN
- China
- Prior art keywords
- msub
- mrow
- mfrac
- aircraft
- mover
- 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.)
- Expired - Fee Related
Links
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/32—Rotors
- B64C27/46—Blades
- B64C27/467—Aerodynamic features
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
- B64U10/00—Type of UAV
- B64U10/10—Rotorcrafts
Abstract
The invention discloses a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability, this method is that the minimum eccentric throw and minimum speed of aircraft are obtained according to resolution of eye and retentivity time of eye；The value of the preliminary geometric parameter for choosing aircraft, Mass Distribution and expected rotating speed；According to the theoretical span for obtaining aerofoil profile steady state stability of spinning stability；The stable solution of kinetic linearity model based on aircraft, obtain the relational expression between aerofoil profile steady state stability and the hovering angle of attack of aircraft；Aerofoil profile steady state stability is chosen in the span of the aerofoil profile steady state stability；Based on the value of the relational expression between the aerofoil profile steady state stability and the hovering angle of attack of aircraft, the aerofoil profile steady state stability chosen, geometric parameter, Mass Distribution and expected rotating speed, the hovering lift of calculating aircraft；Hovering lift after aircraft adjustment is more than self gravitation.This method can obtain the extra lift of aircraft under high-revolving flying method independent of carrying.
Description
Technical field
The invention belongs to unmanned aerial vehicle design technical field, and in particular to a kind of single-blade convolution aircraft based on spinning stability
Lift Optimization Design.
Background technology
Single-blade convolution aircraft is a kind of bionic Aircraft newborn in recent years, uses for reference the hanging principle of samara, during flight
Driving moment is produced by the propulsion plant on wing trailing edge or control arm, whole body is spinned, produces lift.Compared to biography
System Fixed Wing AirVehicle, single-blade convolution aircraft have VTOL, hovering ability, and maneuverability is strong, can fly in various narrow
Small space；Compared to traditional rotor machine, single-blade convolution aircraft mechanical structural is simple, small volume, and naked eyes are difficult to discover during flight, and
Flight noise is small, more suitable for investigation, tracing task.In addition, the characteristics of whole body is in spin states during by flight, auxiliary
With appropriate image acquisition equipment and treatment technology, 360 ° of real time panoramic observations can be carried out, are more beneficial for completing monitoring task.
Because single-blade convolution vehicle technology is still in the starting stage, although current rarely have for vehicle dynamics spy
Property System design method, but because single-blade convolution aircraft is provided simultaneously with the gas of Fixed Wing AirVehicle and rotor craft
Dynamic feature, suffered air force, aerodynamics moment are highly asymmetric, have relative complex dynamics, cause statics to set
Meter criterion is difficult to meet demand, and fixed-wing or rotor design method that can not be traditionally determine parameter.To ensure single-blade convolution
The flight quality of aircraft, reduce control consumption, it is necessary to be self-stabilization by Flight Vehicle Design in Preliminary design, make its
Without steadily hovering can be realized under control state.In addition, this kind of aircraft produces lift by spin at a high speed, maintains hovering, but mistake
High rotating speed can be adversely affected to electronic equipment on measuring cell and machine, it is necessary to which exploitation is a kind of to be obtained by increasing the angle of attack
Extra lift, independent of carrying high-revolving flying method.Therefore, stability Design is optimized for the lift of the vehicle dynamics
Method is indispensable.
The content of the invention
In view of this, the invention provides a kind of single-blade convolution aircraft lift optimization design side based on spinning stability
Method；This method can obtain the extra lift of aircraft under high-revolving flying method independent of carrying.
Realize that specific embodiments of the present invention are as follows：
A kind of single-blade convolution aircraft lift Optimization Design based on spinning stability, is comprised the following steps that：
Step 1, the minimum eccentric throw and minimum speed of aircraft are obtained according to resolution of eye and retentivity time of eye；
Step 2, the value of the preliminary geometric parameter for choosing aircraft, Mass Distribution and expected rotating speed, utilizes the several of aircraft
What parameter and Mass Distribution obtain the expection eccentric throw of aircraft so that it is minimum eccentric that the expection eccentric throw of aircraft is more than it
Away from the expection rotating speed of aircraft is more than its minimum speed；
Step 3, it is theoretical based on foline Momentum mixing, the dynamics nonlinear model of aircraft is obtained, based on the flight
The dynamics nonlinear model of device, using microvariations method, the kinetic linearity model of aircraft is obtained, based on step 2 institute
Geometric parameter, Mass Distribution and the value for being expected rotating speed for the aircraft just chosen, and it is quiet according to the theoretical acquisition aerofoil profile of spinning stability
The span of stability；
Step 4, the stable solution of the kinetic linearity model based on aircraft, obtain aerofoil profile steady state stability and aircraft
Hovering the angle of attack between relational expression；
Step 5, aerofoil profile steady state stability is chosen in the span of the aerofoil profile steady state stability；
Step 6, based on the relational expression between the aerofoil profile steady state stability and the hovering angle of attack of aircraft, the aerofoil profile chosen
Steady state stability, geometric parameter, Mass Distribution and the value for being expected rotating speed, the hovering lift of calculating aircraft；
Step 7, using acquisition aircraft hovering lift compared with self gravitation, if choose aerofoil profile static-stability
The value of degree has traveled through current span, and obtains aircraft hovering lift and be respectively less than self gravitation, repeat step two~six
The repetition for carrying out parameter is chosen；If the value for choosing aerofoil profile steady state stability make it that obtaining aircraft hovering lift is more than self gravitation
When, using the current value for choosing aerofoil profile steady state stability as optimum results, so as to complete the optimization design of aircraft lift.
Further, step 1 detailed process is as follows:
According to human eye vision persistence effect, it is 1/24s, the flight primary election rotating speed r of single-blade convolution aircraft to persist the time0Take
r0>=π/(1/24) ≈ 75.40rad/s, the minimum speed for obtaining aircraft are 75.40rad/s；
The distance definition of pivot to single-blade convolution aircraft main wing root is eccentric distance e, according to the resolving power of human eye
For 2 jiaos points, the minimum eccentric distance e that human eye can be differentiated is obtainedmin=P π/60/180, wherein P are viewing distance.
Further, the detailed process of step 3 is as follows：
2.1 is theoretical based on foline Momentum mixing, and the dynamics nonlinear model for obtaining aircraft is：
Wherein, Vcm=[u, v, w]T, VcmIt is speed under aircraft body system, Ω=[p, q, r]T, Ω is aircraft body
The lower rotational angular velocity of system, Faero=[0, A ,-N]T, FaeroIt is suffered aerodynamic force, M under aircraft body systemaero=[MR,MN,MA]T,
MN、MAAnd MRPitching moment, yawing, rolling moment suffered by respectively whole body, Maero are under aircraft body system
Suffered aerodynamic moment,I is aircraft inertial moment matrix；M is aircraft gross mass, and A and N are respectively to act on
Normal force, axial force in main wing；
2.2 ask for the solution of inearized model, obtain the span of aerofoil profile steady state stability
Angle of attack is considered as low-angle, aircraft is in floating state, and linear velocity is zero along each axis component, around body system Z axis
Rotating speed is fixed as r0, orderWherein, l is integration
Replace variable, IaThe inertia of air involved by main wing rotary motion is represented, c is main chord-length, and ρ is atmospheric density, CM0、CM1Respectively
For aerofoil profile moment coefficient CMConstant term and coefficient of first order, wihInduced velocity during to hover, λihRepresent induced velocity during hovering
Inflow ratio,WithTorque M is represented respectivelyNWith rolling moment MRFitting coefficient normalisation coefft, then
Wherein, η is aerofoil profile steady state stability, φIFor main wing twisting angle, λdFor the inverse of aspect ratio, lwGrown for the main span, e is
Eccentric throw；CL0、CL1Respectively wing section lift coefficient CLConstant term and coefficient of first order；
Using microvariations method, p, q are considered as a small amount of εp、εqAnd ignore εp、εqSecond order it is a small amount of, substitute into rotational power
Equation (7), obtain the kinetic linearity model of aircraft
Wherein
Kinetic linearity solution to model isAccording to the stability of dynamic system
Condition, need a21a12＜ 0, then have
KI(KaKs- 1) ＜ 0
WhereinKIFor inertia parameter,KaFor rotor aerodynamic parameter,
KsFor aerofoil profile steady state stability parameter；
Due to KI＞ 0, then stability condition be reduced to
KaKs＜ 1
After the geometric parameter of aircraft determines, KaIt has been fixed that, now parameter to be designed is Ks；Ensure Ka＞ 0, determines the wing
The span of type steady state stability is
Further, the detailed process of step 4 is as follows：
The kinetic linearity model of the aircraft obtained by step 3, the stable solution of inearized model correspond toWith
Then
Along each axis component of body it is zero according to linear velocity during hovering, obtains the hovering angle of attack of aerofoil profile steady state stability and aircraft
Between relational expression be
Beneficial effect：
(1) The present invention gives the quantification design method for ensureing single-blade convolution aircraft spinning stability, when rotor is pneumatically joined
When number (steady state stability parameter) is preset, can designing steady state stability parameter according to the method for the invention that provide, (rotor is pneumatically joined
Number).
(2) scope of design provided according to the present invention, on the basis of spinning stability is ensured, can relax aerofoil profile steady state stability
Selection, increase balanceable attack angles, to obtain extra lift in hovering, reduce rotating speed needed for hovering.
Brief description of the drawings
Fig. 1 is the flow chart of the present invention；
Fig. 2 is single-blade convolution aircraft top view；
Fig. 3 is aircraft coordinate system and enters flow diagram；
Fig. 4 is become a mandarin for unit foline and stress diagram；
Fig. 5 is main wing steady state stability schematic diagram；
Fig. 6 is aircraft body angular speed；
Fig. 7 is the aircraft construction face angle of attack；
Fig. 8 is aircraft flight track.
Embodiment
The present invention will now be described in detail with reference to the accompanying drawings and examples.
The invention provides a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability, such as Fig. 1 institutes
Show, it is comprised the following steps that：
Step 1：According to human eye vision persistence effect, it is about t=1/24s to persist the time.To ensure stealth effect as design
During premise, then the flight rotating speed r of single-blade convolution aircraft is required0≥π/(1/24)≈75.40rad/s.To avoid swing circle
Approached with human eye data volume frequency acquisition, static images, the weakening stealth effect for causing aircraft to overlap, rotating speed should be avoided to approach
r*=2k π/(1/24) rad/s, k=1,2 ... n.
Eccentric distance e is defined as the distance that pivot (barycenter) arrives single-blade convolution aircraft main wing root, as shown in Figure 2.
When being designed to eccentric throw, it is contemplated that human eye resolution capability under general light conditions is 2 jiaos points, and eccentric throw will be too small to be caused
Aircraft is identified as non-rotary body by human eye, and then loses stealth effect.Assuming that viewing distance is P rice, then human eye can be differentiated
Minimum eccentric distance emin=P π/60/180.So that viewing distance is 100 meters as an example, e can be tried to achievemin=0.029 meter, that is, assume observation
In the case where 100 meters of distances catch aircraft using human eye vision, single-blade convolution aircraft can be realized effectively stealthy inclined person
The heart is away from for 0.029 meter.
Step 2：The preliminary geometric parameter for choosing aircraft, Mass Distribution, aerofoil profile steady state stability and the value for being expected rotating speed,
The expection eccentric throw of aircraft is obtained using the geometric parameter and Mass Distribution of aircraft so that the expection eccentric throw of aircraft is big
In its minimum eccentric throw, the expection rotating speed of aircraft is more than its minimum speed；Protected it must be noted that should try one's best in Preliminary design
Card single-blade convolution aircraft spins around the maximum or minimum principal axis of inertia, and to leave leeway for the adjustment of steady state stability, increase is outstanding
Stop the angle of attack, otherwise be ensure aircraft steadily hovering, it is necessary to select static-stability aerofoil profile with compensate as caused by inertia it is unstable from
Rotation state, the extra hovering angle of attack can not be provided.
Step 3：It is theoretical based on foline Momentum mixing, the dynamics nonlinear model of aircraft is obtained, utilizes microvariations side
Method, obtain the kinetic linearity model of aircraft, based on step 2 at the beginning of choose the geometric parameter of aircraft, Mass Distribution
With the value of expected rotating speed, and according to the theoretical span for obtaining aerofoil profile steady state stability of spinning stability；
2.1 carry out aerodynamic modeling based on foline Momentum mixing theory to aircraft
For accurate simulation single-blade circle round aerodynamic characteristics of vehicle, herein using foline Momentum mixing theoretical description main wing.By
Foline is theoretical to be understood, the inflow velocity of unit foline is
Wherein u, v, w are respectively the speed along body system x, y, z axle, and p, q, r are the angular speed along body system x, y, z axle,
lbeFor unit foline position, wiFor induced velocity, UT、UR、UPIt is the tangential of unit foline, radial direction, normal direction inflow velocity respectively,
As shown in Figure 3.
Because radial direction inflow velocity is smaller to main wing stressing influence, so being ignored to radial direction inflow velocity, must can make
Lift and resistance on unit foline are respectively
Wherein FPFor tip loss correction factor, ρ is atmospheric density, and U is total inflow velocity, and c is main chord-length, CL、CDPoint
Wei not lift, resistance coefficient.
From the figure 3, it may be seen that construction face angle of attack=tan-1(UP/UT), obtain foline angle of attackt=α+φI(lbe), φIBased on ()
Wing twisting angle function, it is fitted, is had according to the foline angle of attack
CLi、CDiRespectively lift, fitting coefficient corresponding to resistance coefficient, i=0,1,2.
Above-mentioned aerodynamic force is changed to body axis system, can be obtained
DN=dL cos α+dD sin α
DA=dD cos α-dL sin α
It is axial force and normal force caused by unit foline, as shown in Figure 4.Above formula is opened up to accumulating along main wing
Point, the axial force A and normal force N suffered by whole body is can obtain, the normal force is main wing caused by aircraft spin
Pulling force.
The coordinate system that becomes a mandarin of the lift and resistance is changed to body axis system, obtained
DN=dL cos α+dD sin α (1)
DA=dD cos α-dL sin α (2)
As the pitching moment M suffered by (1) and (2) obtains whole bodyN, yawing MAWith rolling moment MR：
Wherein e is eccentric throw, lbeFor unit foline position, lwGrown for the main span, CMFor aerofoil profile moment coefficient,For aerofoil profile steady state stability, xcmFor main wing barycenter chordwise location, xaero
For the Aerodynamic force action chordwise location of main wing.
Obtaining single-blade convolution aircraft six-degree-of-freedom dynamic model by (3), (4) and (5) is：
Wherein, Vcm=[u, v, w]T, VcmIt is speed under aircraft body system, Ω=[p, q, r]T, Ω is aircraft body
The lower rotational angular velocity of system, Faero=[0, A ,-N]T, FaeroIt is suffered aerodynamic force, M under aircraft body systemaero=[MR,MN,MA]T,
MN、MAAnd MRPitching moment, yawing, rolling moment suffered by respectively whole body, Maero are under aircraft body system
Suffered aerodynamic moment,I is aircraft inertial moment matrix；M is aircraft gross mass, and G is aircraft in body
The lower suffered gravity of system, A and N are respectively normal force, the axial force for acting on main wing；
2.2 ask for the solution of inearized model, obtain stable spin scope；
Angle of attack is considered as low-angle, aircraft is in floating state, and linear velocity is zero along each axis component, around body system Z axis
Rotating speed is fixed as r0, orderWherein, l is integration
Replace variable, IaThe inertia of air involved by main wing rotary motion is represented, c is main chord-length, and ρ is atmospheric density, CM0、CM1Respectively
For aerofoil profile moment coefficient CMConstant term and coefficient of first order, wihInduced velocity during to hover, λihRepresent induced velocity during hovering
Inflow ratio,WithTorque M is represented respectivelyNWith rolling moment MRFitting coefficient normalisation coefft, then
Wherein, η is aerofoil profile steady state stability, φIFor main wing twisting angle, λdFor the inverse of aspect ratio, lwGrown for the main span, e is
Eccentric throw；CL0、CL1Respectively wing section lift coefficient CLConstant term and coefficient of first order；
Using microvariations method, p, q are considered as a small amount of εp、εqAnd ignore εp、εqSecond order it is a small amount of, substitute into rotational power
Equation (7), obtain the kinetic linearity model of aircraft
Wherein
Wherein, a12、a21、a22For element in matrix Α, b1、b2For element in matrix B；
Kinetic linearity solution to model isAccording to the stability of dynamic system
Condition, need a21a12＜ 0, then have
KI(KaKs- 1) ＜ 0
WhereinKIFor inertia parameter,KaFor rotor aerodynamic parameter,
KsFor aerofoil profile steady state stability parameter；
Due to KI＞ 0, then stability condition be reduced to
KaKs＜ 1
After the geometric parameter of aircraft determines, KaIt has been fixed that, now parameter to be designed is Ks；Ensure Ka＞ 0, determines the wing
The span of type steady state stability isAs shown in Figure 5.
Step 4：The stable solution of kinetic linearity model based on aircraft, obtain aerofoil profile steady state stability and aircraft
Hovering the angle of attack between relational expression；
According to the kinetic linearity model stability solution of aircraft, increasable hovering after change aerofoil profile steady state stability is calculated
The angle of attack, so as to obtain extra lift.
By the kinetic linearity model of the obtained single-blade convolution aircraft in upper step, orderStable solution, which can be obtained, is
Linear velocity is zero along each axis component of body during due to hovering, ignores induced velocity influence and nonlinear factor CM,
The relational expression that can be obtained between aerofoil profile steady state stability and the hovering angle of attack of aircraft is
Step 5, aerofoil profile steady state stability is chosen in the span of the aerofoil profile steady state stability；
Step 6, relational expression, the selection between the hovering angle of attack based on the aerofoil profile steady state stability and aircraft
The value of the aerofoil profile steady state stability and preliminary geometric parameter for choosing aircraft, Mass Distribution and expected rotating speed, calculating aircraft
Hovering lift；
Step 7, using acquisition aircraft hovering lift compared with self gravitation, if choose aerofoil profile static-stability
The value of degree has traveled through current span, and obtains aircraft hovering lift and be respectively less than self gravitation, repeat step two~six
The repetition for carrying out parameter is chosen；If the value for choosing aerofoil profile steady state stability make it that obtaining aircraft hovering lift is more than self gravitation
When, using the current value for choosing aerofoil profile steady state stability as optimum results, so as to complete the optimization design of aircraft lift.
Stability Design standard is had been presented in upper step, aerofoil profile steady state stability is carried out again according to the scope that the standard provides
Design, can draw the new hovering angle of attack.Bring the new hovering angle of attack into lift calculation procedure, judge whether to meet lift demand
And spinning stability demand.If be unsatisfactory for, then steady state stability is chosen from new, above-mentioned flow is repeated, until meeting index request.
Embodiment
By taking certain pre-designed single-blade convolution aircraft as an example, illustrate the specific implementation method of the present invention.Single-blade convolution flight
Device Design Fundamentals parameter is as shown in table 1.
Table 1.
| Parameter | Symbol | Value | Unit |
| Atmospheric density | ρ | 1.205 | kg/m3 |
| Rotary inertia | Ix | 0.0018 | kg·m2 |
| Iy | 0.0029 | kg·m2 | |
| Iz | 0.0047 | kg·m2 | |
| Gross mass | m | 0.22 | kg |
| Eccentric throw | e | 0.0362 | m |
| The span | lw | 0.3 | m |
| Chord length | c | 0.08 | m |
| Preset rotation speed | r0 | -105 | rad·s-1 |
| Lift coefficient | CL1 | 0.1018 | deg-1 |
| CL0 | 0.2973 | ||
| Moment coefficient | CM1 | 0.0528 | deg-1 |
The method provided according to step 2, obtain in the case where being added without aerofoil profile steady state stability, this group of design parameter institute
The pulling force that can be provided is N=2.0777N, less than gravity G=2.1560N suffered by aircraft.Therefore step 3 is performed.
I can be calculated according to the parameter of table 1a=ρ cl4=0.00078084,
According to the stability condition provided in step 3, have
Due to KsMiddle undetermined parameter is steady state stability η, can be obtained
η ＜ 509.4990
In the scope of design provided, it is quiet unstable by Airfoil Design, chooses η=0.3.Calculating in step 4
Formula, obtain
Lift is calculated using the increased angle of attack of steady state stability institute after adjustment, N=2.1540N can be obtained, suffered by aircraft
The difference of gravity is less than 0.1%.Simulation result is as shown in figs 6-8.
In summary, presently preferred embodiments of the present invention is these are only, is not intended to limit the scope of the present invention.
Within the spirit and principles of the invention, any modification, equivalent substitution and improvements made etc., it should be included in the present invention's
Within protection domain.
Claims (4)
1. a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability, it is characterised in that specific steps are such as
Under：
Step 1, the minimum eccentric throw and minimum speed of aircraft are obtained according to resolution of eye and retentivity time of eye；
Step 2, the value of the preliminary geometric parameter for choosing aircraft, Mass Distribution and expected rotating speed, utilizes the geometric parameters of aircraft
Number and Mass Distribution obtain the expection eccentric throw of aircraft so that the expection eccentric throw of aircraft is more than its minimum eccentric throw, flies
The expection rotating speed of row device is more than its minimum speed；
Step 3, it is theoretical based on foline Momentum mixing, the dynamics nonlinear model of aircraft is obtained, based on the aircraft
Dynamics nonlinear model, using microvariations method, the kinetic linearity model of aircraft is obtained, based on the primary election of step 2 institute
Geometric parameter, Mass Distribution and the value for being expected rotating speed of the aircraft taken, and obtain aerofoil profile static-stability according to spinning stability is theoretical
The span of degree；
Step 4, the stable solution of the kinetic linearity model based on aircraft, acquisition aerofoil profile steady state stability and aircraft hang
Stop the relational expression between the angle of attack；
Step 5, aerofoil profile steady state stability is chosen in the span of the aerofoil profile steady state stability；
Step 6, it is quiet steady based on the relational expression between the aerofoil profile steady state stability and the hovering angle of attack of aircraft, the aerofoil profile chosen
Fixed degree, geometric parameter, Mass Distribution and the value for being expected rotating speed, the hovering lift of calculating aircraft；
Step 7, using acquisition aircraft hovering lift compared with self gravitation, if choosing aerofoil profile steady state stability
Value has traveled through current span, and obtains aircraft hovering lift and be respectively less than self gravitation, repeat step two~six；If choosing
When taking the value of aerofoil profile steady state stability make it that acquisition aircraft hovering lift is more than self gravitation, aerofoil profile steady state stability will be currently chosen
Value as optimum results, so as to complete the optimization design of aircraft lift.
2. a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability as claimed in claim 1, its feature
It is, step 1 detailed process is as follows:
According to human eye vision persistence effect, it is 1/24s, the flight primary election rotating speed r of single-blade convolution aircraft to persist the time0Take r0≥
π/(1/24) ≈ 75.40rad/s, the minimum speed for obtaining aircraft are 75.40rad/s；
The distance definition of pivot to single-blade convolution aircraft main wing root is eccentric distance e, is 2 jiaos according to the resolving power of human eye
Point, obtain the minimum eccentric distance e that human eye can be differentiatedmin=P π/60/180, wherein P are viewing distance.
3. a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability as claimed in claim 1, its feature
It is, the detailed process of step 3 is as follows：
2.1 is theoretical based on foline Momentum mixing, and the dynamics nonlinear model for obtaining aircraft is：
<mrow>
<mi>m</mi>
<msub>
<mover>
<mi>V</mi>
<mo>&CenterDot;</mo>
</mover>
<mrow>
<mi>c</mi>
<mi>m</mi>
</mrow>
</msub>
<mo>=</mo>
<msub>
<mi>F</mi>
<mrow>
<mi>a</mi>
<mi>e</mi>
<mi>r</mi>
<mi>o</mi>
</mrow>
</msub>
<mo>+</mo>
<mi>m</mi>
<mi>g</mi>
<mo>-</mo>
<mi>m</mi>
<mi>&Omega;</mi>
<mo>&times;</mo>
<msub>
<mi>V</mi>
<mrow>
<mi>c</mi>
<mi>m</mi>
</mrow>
</msub>
<mo>-</mo>
<mo>-</mo>
<mo>-</mo>
<mrow>
<mo>(</mo>
<mn>6</mn>
<mo>)</mo>
</mrow>
</mrow>
<mrow>
<mi>I</mi>
<mover>
<mi>&Omega;</mi>
<mo>&CenterDot;</mo>
</mover>
<mo>=</mo>
<msub>
<mi>M</mi>
<mrow>
<mi>a</mi>
<mi>e</mi>
<mi>r</mi>
<mi>o</mi>
</mrow>
</msub>
<mo>-</mo>
<mi>&Omega;</mi>
<mo>&times;</mo>
<mrow>
<mo>(</mo>
<mi>I</mi>
<mi>&Omega;</mi>
<mo>)</mo>
</mrow>
<mo>-</mo>
<mo>-</mo>
<mo>-</mo>
<mrow>
<mo>(</mo>
<mn>7</mn>
<mo>)</mo>
</mrow>
</mrow>
Wherein, Vcm=[u, v, w]T, VcmIt is speed under aircraft body system, Ω=[p, q, r]T, Ω is under aircraft body system
Rotational angular velocity, Faero=[0, A ,-N]T, FaeroIt is suffered aerodynamic force, M under aircraft body systemaero=[MR,MN,MA]T, MN、
MAAnd MRPitching moment, yawing, rolling moment suffered by respectively whole body, MaeroIt is suffered under aircraft body system
Aerodynamic moment,I is aircraft inertial moment matrix；M is aircraft gross mass, and A and N are respectively to act on master
Normal force, the axial force of the wing；
2.2 ask for the solution of inearized model, obtain the span of aerofoil profile steady state stability
Angle of attack is considered as low-angle, aircraft is in floating state, and linear velocity is zero along each axis component, and flight primary election rotating speed is consolidated
It is set to r0, make Ia=ρ cl4,Wherein, l is replaced for integration and become
Amount, IaThe inertia of air involved by main wing rotary motion is represented, c is main chord-length, and ρ is atmospheric density, CM0、CM1Respectively aerofoil profile
Moment coefficient CMConstant term and coefficient of first order, wihInduced velocity during to hover, λihInduced velocity inflow ratio during hovering is represented,WithPitching moment M is represented respectivelyNWith rolling moment MRFitting coefficient normalisation coefft, then
<mrow>
<msub>
<mi>M</mi>
<mi>N</mi>
</msub>
<mo>=</mo>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>8</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>8</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<mfrac>
<mi>q</mi>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>8</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
</mrow>
<mfenced open = "" close = "">
<mtable>
<mtr>
<mtd>
<mrow>
<msub>
<mi>M</mi>
<mi>R</mi>
</msub>
<mo>=</mo>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<mfrac>
<mi>q</mi>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
</mrow>
</mtd>
</mtr>
<mtr>
<mtd>
<mrow>
<mo>+</mo>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<msub>
<mi>&eta;&lambda;</mi>
<mi>d</mi>
</msub>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<mfrac>
<mi>q</mi>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
</mrow>
</mtd>
</mtr>
</mtable>
</mfenced>
Wherein, η is aerofoil profile steady state stability, φIFor main wing twisting angle, λdFor the inverse of aspect ratio, lwGrown for the main span, e is bias
Away from；CL0、CL1Respectively wing section lift coefficient CLConstant term and coefficient of first order；
Using microvariations method, p, q are considered as a small amount of εp、εqAnd ignore εp、εqSecond order it is a small amount of, substitute into rotational power equation
(7) the kinetic linearity model of aircraft, is obtained
<mrow>
<mover>
<mi>&epsiv;</mi>
<mo>&CenterDot;</mo>
</mover>
<mo>=</mo>
<mi>A</mi>
<mi>&epsiv;</mi>
<mo>+</mo>
<mi>B</mi>
</mrow>
Wherein
<mrow>
<mi>&epsiv;</mi>
<mo>=</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>&epsiv;</mi>
<mi>p</mi>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>&epsiv;</mi>
<mi>q</mi>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>,</mo>
<mi>A</mi>
<mo>=</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<mn>0</mn>
</mtd>
<mtd>
<msub>
<mi>a</mi>
<mn>12</mn>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>a</mi>
<mn>21</mn>
</msub>
</mtd>
<mtd>
<msub>
<mi>a</mi>
<mn>22</mn>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
<mo>,</mo>
<mi>B</mi>
<mo>=</mo>
<mfenced open = "[" close = "]">
<mtable>
<mtr>
<mtd>
<msub>
<mi>b</mi>
<mn>1</mn>
</msub>
</mtd>
</mtr>
<mtr>
<mtd>
<msub>
<mi>b</mi>
<mn>2</mn>
</msub>
</mtd>
</mtr>
</mtable>
</mfenced>
</mrow>
<mrow>
<msub>
<mi>a</mi>
<mn>12</mn>
</msub>
<mo>=</mo>
<mfrac>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mfrac>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>&eta;&lambda;</mi>
<mi>d</mi>
</msub>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>-</mo>
<mfrac>
<mrow>
<msub>
<mi>I</mi>
<mi>z</mi>
</msub>
<mo>-</mo>
<msub>
<mi>I</mi>
<mi>y</mi>
</msub>
</mrow>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mfrac>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<mo>,</mo>
</mrow>
<mrow>
<msub>
<mi>a</mi>
<mn>21</mn>
</msub>
<mo>=</mo>
<mfrac>
<mrow>
<msub>
<mi>I</mi>
<mi>z</mi>
</msub>
<mo>-</mo>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mrow>
<msub>
<mi>I</mi>
<mi>y</mi>
</msub>
</mfrac>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<mo>,</mo>
</mrow>
<mrow>
<msub>
<mi>a</mi>
<mn>22</mn>
</msub>
<mo>=</mo>
<mfrac>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<mrow>
<mn>8</mn>
<msub>
<mi>I</mi>
<mi>y</mi>
</msub>
</mrow>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>,</mo>
</mrow>
<mrow>
<msub>
<mi>b</mi>
<mn>1</mn>
</msub>
<mo>=</mo>
<mfrac>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mfrac>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>0</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>+</mo>
<mfrac>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mfrac>
<msub>
<mi>&eta;&lambda;</mi>
<mi>d</mi>
</msub>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>,</mo>
</mrow>
<mrow>
<msub>
<mi>b</mi>
<mn>2</mn>
</msub>
<mo>=</mo>
<mfrac>
<msub>
<mi>I</mi>
<mi>a</mi>
</msub>
<msub>
<mi>I</mi>
<mi>y</mi>
</msub>
</mfrac>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>8</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>8</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<msubsup>
<mi>r</mi>
<mn>0</mn>
<mn>2</mn>
</msubsup>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>;</mo>
</mrow>
Kinetic linearity solution to model isAccording to the stability condition of dynamic system,
Need a21a12＜ 0, then have
KI(KaKs- 1) ＜ 0
WhereinKIFor inertia parameter,KaFor rotor aerodynamic parameter,
KsFor aerofoil profile steady state stability parameter；
Due to KI＞ 0, then stability condition be reduced to
KaKs＜ 1
After the geometric parameter of aircraft determines, KaIt has been fixed that, now parameter to be designed is Ks；Ensure Ka＞ 0, determine that aerofoil profile is quiet
The span of stability is
<mrow>
<mi>&eta;</mi>
<mo><</mo>
<mfrac>
<mrow>
<mfrac>
<mn>6</mn>
<msub>
<mi>K</mi>
<mi>a</mi>
</msub>
</mfrac>
<mo>-</mo>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
</mrow>
<mrow>
<msub>
<mi>&lambda;</mi>
<mi>d</mi>
</msub>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
</mrow>
</mfrac>
<mo>.</mo>
</mrow>
4. a kind of single-blade convolution aircraft lift Optimization Design based on spinning stability as claimed in claim 3, its feature
It is, the detailed process of step 4 is as follows：
The kinetic linearity model of the aircraft obtained by step 3, the stable solution of inearized model correspond toWithThen
<mrow>
<msub>
<mover>
<mi>&epsiv;</mi>
<mo>~</mo>
</mover>
<mi>q</mi>
</msub>
<mo>=</mo>
<mo>-</mo>
<mfrac>
<msub>
<mi>b</mi>
<mn>1</mn>
</msub>
<msub>
<mi>a</mi>
<mn>12</mn>
</msub>
</mfrac>
</mrow>
<mrow>
<msub>
<mover>
<mi>&epsiv;</mi>
<mo>~</mo>
</mover>
<mi>p</mi>
</msub>
<mo>=</mo>
<mfrac>
<mrow>
<msub>
<mi>a</mi>
<mn>22</mn>
</msub>
<msub>
<mi>b</mi>
<mn>1</mn>
</msub>
</mrow>
<mrow>
<msub>
<mi>a</mi>
<mn>12</mn>
</msub>
<msub>
<mi>a</mi>
<mn>21</mn>
</msub>
</mrow>
</mfrac>
<mo>-</mo>
<mfrac>
<msub>
<mi>b</mi>
<mn>2</mn>
</msub>
<msub>
<mi>a</mi>
<mn>21</mn>
</msub>
</mfrac>
</mrow>
Along each axis component of body it is zero according to linear velocity during hovering, obtains between aerofoil profile steady state stability and the hovering angle of attack of aircraft
Relational expression be
<mrow>
<mover>
<mi>&alpha;</mi>
<mo>~</mo>
</mover>
<mo>=</mo>
<mfrac>
<msub>
<mover>
<mi>&epsiv;</mi>
<mo>~</mo>
</mover>
<mi>q</mi>
</msub>
<msub>
<mi>r</mi>
<mn>0</mn>
</msub>
</mfrac>
<mo>=</mo>
<mo>-</mo>
<mfrac>
<mrow>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>0</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
<mo>+</mo>
<msub>
<mi>&eta;&lambda;</mi>
<mi>d</mi>
</msub>
<mo>(</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<msub>
<mi>&phi;</mi>
<mi>I</mi>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>4</mn>
</mfrac>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>0</mn>
</mrow>
</msub>
<msub>
<mi>&lambda;</mi>
<mrow>
<mi>i</mi>
<mi>h</mi>
</mrow>
</msub>
<mo>)</mo>
</mrow>
<mrow>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mover>
<mi>C</mi>
<mo>&OverBar;</mo>
</mover>
<mrow>
<mi>M</mi>
<mn>1</mn>
</mrow>
</msub>
<mo>+</mo>
<mfrac>
<mn>1</mn>
<mn>6</mn>
</mfrac>
<msub>
<mi>&eta;&lambda;</mi>
<mi>d</mi>
</msub>
<msub>
<mi>C</mi>
<mrow>
<mi>L</mi>
<mn>1</mn>
</mrow>
</msub>
<mo>-</mo>
<mfrac>
<mrow>
<msub>
<mi>I</mi>
<mi>z</mi>
</msub>
<mo>-</mo>
<msub>
<mi>I</mi>
<mi>y</mi>
</msub>
</mrow>
<msub>
<mi>I</mi>
<mi>x</mi>
</msub>
</mfrac>
</mrow>
</mfrac>
<msubsup>
<mo>|</mo>
<mi>e</mi>
<mrow>
<mi>e</mi>
<mo>+</mo>
<msub>
<mi>l</mi>
<mi>w</mi>
</msub>
</mrow>
</msubsup>
<mo>.</mo>
</mrow>
3
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610329445.3A CN105905296B (en) | 2016-05-18 | 2016-05-18 | Single-blade convolution aircraft lift Optimization Design based on spinning stability |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201610329445.3A CN105905296B (en) | 2016-05-18 | 2016-05-18 | Single-blade convolution aircraft lift Optimization Design based on spinning stability |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN105905296A CN105905296A (en) | 2016-08-31 |
| CN105905296B true CN105905296B (en) | 2017-12-15 |
Family
ID=56749395
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201610329445.3A Expired - Fee Related CN105905296B (en) | 2016-05-18 | 2016-05-18 | Single-blade convolution aircraft lift Optimization Design based on spinning stability |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN105905296B (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106777495B (en) * | 2016-11-22 | 2020-01-17 | 西北工业大学 | Method for controlling rotor wing profile pitching moment through local parametric modification |
| CN110422326A (en) * | 2019-04-30 | 2019-11-08 | 李泽波 | A kind of course of new aircraft and its control method |
| CN112224386A (en) * | 2020-10-20 | 2021-01-15 | 南京航空航天大学 | Single-wing self-rotating landing device |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US7059931B2 (en) * | 2003-05-27 | 2006-06-13 | Veratech Aero-Rpv Corporation | Reduced visibility rotorcraft and method of controlling flight of reduced visibility rotorcraft |
| US20060183398A1 (en) * | 2004-11-24 | 2006-08-17 | Barnes Theodore D | Single-winged auto rotating toy glider |
| US8366055B2 (en) * | 2009-06-19 | 2013-02-05 | University Of Maryland | Controllable miniature mono-wing aircraft |
| US9199718B2 (en) * | 2011-02-12 | 2015-12-01 | Shaun P. Fogarty | Auto-rotating device |
| CN103693195B (en) * | 2013-12-17 | 2015-08-12 | 南京航空航天大学 | A kind of minute vehicle |
-
2016
- 2016-05-18 CN CN201610329445.3A patent/CN105905296B/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| CN105905296A (en) | 2016-08-31 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106249745B (en) | The control method of four axis unmanned planes | |
| Prouty | Helicopter Aerodynamics Volume I | |
| CN206202665U (en) | Propeller, Power Component and aircraft | |
| CN105905296B (en) | Single-blade convolution aircraft lift Optimization Design based on spinning stability | |
| CN107117300B (en) | Unmanned vehicle based on coaxial more rotor pose adjustments | |
| CN109543271B (en) | Trimming control design method applied to coaxial high-speed helicopter with thrust propeller | |
| CN105468814B (en) | A kind of helicopter flight characteristic computing method considering full machine aerodynamic interference | |
| CN110162933B (en) | Coaxial multi-rotor simulation method and system | |
| CN104950901A (en) | Nonlinear robust control method with finite-time convergence capacity for unmanned helicopter attitude error | |
| CN105488295A (en) | Unmanned aerial vehicle modeling system taking wind field disturbances into consideration | |
| CN109612676A (en) | Aerodynamic parameter reverse calculation algorithms based on test flight data | |
| CN109063391A (en) | Dynamic derivative under rotating condition calculates detection method and dynamic derivative wind tunnel test methods | |
| CN116187199A (en) | Unsteady aerodynamic modeling method integrating intelligent technology | |
| CN107521680A (en) | A kind of active wind resistance coaxal helicopter and active wind resistance attitude control method | |
| Rushdi et al. | Simulation of the transition phase for an optimally-controlled tethered vtol rigid aircraft for airbornewind energy generation | |
| Li et al. | Experimental investigation on aerodynamics of nonplanar rotor pairs in a multi-rotor UAV | |
| Rimkus et al. | An application of the autogyro theory to airborne wind energy extraction | |
| Low et al. | Analysis of wing twist effects on hover flight dynamics of a single rotor aerial craft | |
| CN111581795B (en) | Angular motion complex analysis method of rotary aircraft controlled by a pair of duck rudders | |
| Ma et al. | Conceptual Design of a VTOL Box-wing UAV with Rotatable Duct-fans | |
| Runco et al. | Experimental studies on a mesoscale cycloidal rotor in hover | |
| CN112329155A (en) | Simulation method and device of cross dual-rotor unmanned helicopter | |
| Elango et al. | Periodic Flight of Rotorcraft within Thermal Columns | |
| Lind et al. | Reynolds number effects on airfoils in reverse flow | |
| Aliabadi et al. | Experimental study of flapping-wing aerodynamic coefficients and landing performance estimation |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| CB03 | Change of inventor or designer information |
Inventor after: Kang Shen Inventor after: Wang Jianan Inventor after: Dan Jiayuan Inventor before: Wang Jianan Inventor before: Kang Shen Inventor before: Shan Jiayuan |
|
| CB03 | Change of inventor or designer information | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20171215 Termination date: 20190518 |
|
| CF01 | Termination of patent right due to non-payment of annual fee |