CN109018422B - Tilting corridor calculation method of fixed-rotation-speed and periodic-pitch tilting four-rotor aircraft - Google Patents

Tilting corridor calculation method of fixed-rotation-speed and periodic-pitch tilting four-rotor aircraft Download PDF

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CN109018422B
CN109018422B CN201810750002.0A CN201810750002A CN109018422B CN 109018422 B CN109018422 B CN 109018422B CN 201810750002 A CN201810750002 A CN 201810750002A CN 109018422 B CN109018422 B CN 109018422B
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潘浙平
陈仁良
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Nanjing University of Aeronautics and Astronautics
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Abstract

The invention discloses a method for calculating a tilting corridor of a fixed-speed tilting quad-rotor aircraft with periodic variable pitch, which comprises the following steps of: establishing a flight mechanics model of each part of the aircraft, converting all forces and moments into a body axis system, and establishing a six-degree-of-freedom balance equation; determining a tilting operation mode; solving the upper and lower boundaries of a pitch angle by taking the lift force characteristic of the wing as a limiting factor; carrying out optimal balancing solution to obtain the manipulated variable and the attitude angle under different state conditions; finding a minimum speed limit boundary and a high speed limit boundary corresponding to the wing lift characteristic; and solving the required power at different state points to obtain the available power limit high-speed boundary. The invention is based on a flight mechanics model, considers parameters of various aspects of aircraft design, and reflects the specific attributes of each feasible flight point. The method is also suitable for calculating the tilting corridor of the tilting quadrotor aircraft in different operation modes.

Description

Tilting corridor calculation method of fixed-rotation-speed and periodic-pitch tilting four-rotor aircraft
Technical Field
The invention discloses a method for calculating a tilting corridor of a tilting quadrotor (QTR for short) with fixed rotating speed and periodic variable pitch control, belonging to the technical field of flight mechanics and control of the tilting quadrotor.
Background
Since the concept of QTR proposed by the Bell helicopter company in America at the beginning of this century, the QTR is still in an exploration stage at home and abroad. The configuration is shown in figure 1, and the aircraft is provided with a front auxiliary wing and a rear auxiliary wing, wherein a tiltable rotor nacelle is arranged at two ends of the wing.
The QTR vertically takes off and lands in a helicopter mode, flies at high speed in a fixed-wing airplane mode, and can tilt to enter a tilting transition mode through a nacelle. Therefore, the QTR has the advantages of vertical take-off and landing in a helicopter mode, fixed-point hovering, remote transportation in an airplane mode and high-speed cruising, a pair of wings and a pair of rotors are added relative to the tilting dual rotors, maneuverability, carrying capacity and gravity center adjustable capacity are greatly enhanced, and the QTR has very important research value in military and civil transportation.
There are few published literature publications relating to QTR in foreign countries. The research aiming at the QTR in China is only in the starting stage. In general, research on QTR at home and abroad is still in an exploration phase.
The tilting transition flight is the most critical and dangerous flight state of QTR because the whole aircraft is in a variant and variable speed state. The size and the width of the tilting transition corridor are key factors for evaluating the tilting difficulty and the safety of the aircraft, and are related to the general parameters of the QTR, the pneumatic layout, the operation mode and the like.
At present, the tilting transition corridor of the tilting dual-rotor aircraft (such as XV-15) is researched, and little research report is made on the QTR tilting transition corridor. The gravity centers of the tilting dual rotors are near the aerodynamic resultant force center of the rotor/wing, and the tilting corridor is calculated based on the particle model and has certain rationality. For QTR, first, the range of variation of the center of gravity is large, and the aerodynamic forces of the four rotors/wings generate relatively large moments, thereby affecting the attitude. Therefore, the particle model is not suitable for QTR, and a flight dynamics model needs to be established. Secondly, QTR control mode is selected abundantly, and different control modes can obtain different tilting corridors.
Therefore, based on the background analysis, the calculation method of the QTR transition corridor is of great significance to the overall design, component design, flight quality, control strategy and control system design of the QTR.
Disclosure of Invention
The invention aims to provide a calculation method for a tilting corridor of a tilting four-rotor aircraft with fixed rotating speed and periodic variable pitch, so as to overcome the defect that a proper method cannot be adopted to calculate the QTR tilting corridor in the prior art. The invention is based on a flight mechanics model, considers parameters of various aspects of aircraft design, and reflects the specific attributes of each feasible flight point. The method is also suitable for calculating the tilting corridor of the tilting quadrotor aircraft in different operation modes.
In order to achieve the purpose, the technical scheme adopted by the invention is as follows:
the invention provides a method for calculating a tilting corridor of a fixed-speed tilting quad-rotor aircraft with periodic variable pitch, which comprises the following steps of:
1) establishing a flight mechanics model of each part of the aircraft, converting all forces and moments into a body axis system, and establishing a six-degree-of-freedom balance equation;
2) selecting a tilting operation mode;
3) solving the upper and lower boundaries of a pitch angle by taking the lift force characteristic of the wing as a limiting factor; carrying out optimal balancing solution to obtain the manipulated variable and the attitude angle under different state conditions; finding a minimum speed limit boundary and a high speed limit boundary corresponding to the wing lift characteristic;
4) and solving the required power at different state points to obtain the available power limit high-speed boundary.
Preferably, the flight mechanics model in step 1) includes: rotor model, wing model, rotor nacelle model and fuselage model.
Preferably, the step 1) further comprises:
11) the rotor model building method comprises the following steps:
a. inflow model
Induction speed
Figure BDA0001725320280000021
Expressed as rotor spanwise position in the form of a first order Fourier series
Figure BDA0001725320280000022
And azimuth ψ:
Figure BDA0001725320280000023
in the formula, λ0，λ1c，λ1sThe two components are respectively equal inflow, first-order longitudinal inflow component and first-order transverse inflow component of the rotor wing;
Figure BDA0001725320280000024
in the formula, the matrix M is an inertia matrix and reflects the time delay of inflow; the V matrix is a mass flow matrix;
Figure BDA0001725320280000025
the matrix is a static relation matrix between the disturbance induction speed and the disturbance pneumatic load; cT，CL，CMAre respectively rotor liftsForce coefficient, aerodynamic roll torque and aerodynamic pitch torque;
considering the interference between the front rotor and the rear rotor, introducing the interference factor of the front rotor and the rear rotor, and the inflow ratio lambda of the front rotor and the rear rotorFAnd λRThe calculation formula is as follows:
Figure BDA0001725320280000026
in formula (II), lambda'F0And λ'R0The free incoming flow ratio of the front rotor wing and the rear rotor wing is respectively; v isF0V and vR0The induced speeds of the front and rear rotors are respectively; dFRFAnd dFFRInterference factors of the rear-to-front rotor and the front-to-rear rotor are respectively;
b. waving motion model
The dynamic equation of the plane of the trajectory of the rotor disk under the rotor shafting is expressed as follows:
Figure BDA0001725320280000031
in the formula, a0、a1s、b1sRespectively is a rotor flapping taper angle, a back chamfer angle and a side inclination angle;
Figure BDA0001725320280000032
is a rotor damping matrix;
Figure BDA0001725320280000033
is a rotor stiffness matrix;
Figure BDA0001725320280000034
is a rotor excitation vector;
c. rotor aerodynamic force
According to the rotor wing element momentum theory, obtaining the tension T of a rotor wing under a hub wind axis system, the backward force H of the rotor wing, the lateral force Y of the rotor wing, the pitching moment M of the rotor wing, the rolling moment L of the rotor wing and the counter torque Q of the rotor wing; converting the force and moment calculated under the rotor wind axis system to the body axis system;
12) the establishment of the wing model specifically comprises the following steps:
the aerodynamic force of the wing-nacelle is divided into two parts: one part is wing aerodynamic force influenced by rotor wake, and the other part is wing aerodynamic force not influenced by rotor wake; area S of wing in slipstream partwssAnd an area S in the free stream partwfsThe calculation formulas of (A) and (B) are respectively as follows:
Figure BDA0001725320280000035
in the formula, thetanIs a tilt angle; sWIs the wing area; sssmax＝2ηssRc，ηssIs a rotor wing slipstream correction factor, R is the rotor wing radius, and c is the wing average aerodynamic chord length; 1.386, 3.114; mu is the rotor wing advancing ratio; mu.smaxThe maximum advance ratio of the rotor tail deviating out of the wing;
the wing lift and drag are respectively:
Figure BDA0001725320280000036
in the formula, qwDynamic pressure of incoming flow at the wing; cLw，CDwThe lift and drag coefficients of the wing are respectively;
13) the establishment of the rotor wing nacelle model specifically comprises the following steps:
rotor nacelle is changeing the in-process, and the area that faces the wind changes, and aerodynamic resistance is the function of nacelle angle of attack, ignores other aerodynamic force and moment:
Dn＝4qnCDn[SnTopcos(αn)+SnSidesin(αn)] (7)
in the formula, qnThe pressure of the incoming flow at the nacelle; cDnIs the nacelle drag coefficient; snTopThe area of the top of the nacelle; snSideIs the nacelle side area; alpha is alphanIs the nacelle angle of attack;
14) the establishment of the fuselage model specifically comprises the following steps:
obtaining a group of airframe aerodynamic coefficients expressed in a dimensionless form, namely resistance coefficients C through a wind tunnel testDbCoefficient of lift CLbCoefficient of lateral force CSbRolling moment coefficient CMxbPitching moment coefficient CMybAnd yaw moment coefficient CMzb(ii) a The interference of a rotor wing and a wing on a fuselage is not considered;
the fuselage aerodynamic forces and moments are expressed as:
Figure BDA0001725320280000041
in the formula IbIs a characteristic length of the fuselage, AbIs a characteristic area, qbThe pressure of the incoming flow at the fuselage.
Preferably, the step 1) specifically further comprises: establishing a hybrid control model through an accelerator/collective pitch rod deltacolLongitudinal push rod deltalonTransverse push rod deltalatPedal deltapedThe vertical channel, the longitudinal channel, the transverse channel and the course are controlled in a mixed mode; the controlled variables are the total pitch, the longitudinal and transverse cyclic variable pitch, the four aileron deflection angles or the subsets thereof of the four ailerons, and the number of the controlled variables is set as n according to the specific operation modecThen, there are:
Figure BDA0001725320280000042
wherein u is a 4 × 1-dimensional manipulated variable vector; c is ncX 1 control variable vector; g is ncA gain matrix of x 4-dimensional manipulated variable to controlled variable;
the W weight distribution matrix is a function along with the tilting angle, and the weight changes according to a sine rule; using the controlled quantity as the total distance of four rotors
Figure BDA0001725320280000045
And four aileron deflection angles deltaflap_iA total of 8 control variables are taken as examples, i is 1,2,3, 4; then there are:
Figure BDA0001725320280000043
preferably, the step 1) specifically further comprises: substituting the aerodynamic force and the moment obtained by the aerodynamic models of all parts of the four-rotor aircraft into an organism motion equation to obtain a nonlinear flight dynamics model expressed in the form of a first-order differential equation:
Figure BDA0001725320280000044
where y represents the total state quantity, u is a 4 × 1-dimensional manipulated variable vector, and t is time.
Preferably, the tilt operation mode selected in step 2) is specifically a fixed-speed periodic pitch operation: the forward force of the front flight of the aircraft is generated by longitudinal periodic variable pitch, and is provided by tilting of the nacelle after the maximum limit of the periodic variable pitch is reached; gradually adding ailerons for operation; pitch attitude can also provide a portion of the forward force; the pitching attitude of the airframe is mainly maintained by the collective pitch differential of the front and rear rotors.
Preferably, the step 3) is specifically: wing angle of attack exceeds wing stall angle of attack alphaw_stallWhen the airplane flies, the airplane wings are considered to be stalled and can not fly normally; when the wing angle of attack is less than zero lift angle of attack alphaw_zeroWhen the wing is in use, the wing can not generate lift force and becomes a load; the pitch angle is determined by
Figure BDA0001725320280000053
The range is as follows:
Figure BDA0001725320280000051
in the formula, alphafw0，αrw0Respectively the initial mounting angles of the front and rear wings;
selecting a group of target pitch angles and tilt angles to be calculated, and respectively carrying out balancing calculation on each target pitch angle and each target tilt angle; carrying out optimization calculation by taking the trim pitch angle as an optimization target to obtain an optimized longitudinal periodic variable pitch; then, determining the lowest speed and the highest speed in the state through interpolation;
and (4) putting all the target attitudes together after completing calculation, and taking the union of all the feasible flight domains to be the limit boundary of the lift characteristic of the wing.
Preferably, the step 4) is specifically: in the process of trim calculation, the torque of the rotor wing is extracted to calculate the required power, and the required power of the whole machine is as follows:
Figure BDA0001725320280000052
in the formula, P is required power; qiThe reaction torque of the ith auxiliary rotor wing;
then only the check is needed: p is less than or equal to Pky(ii) a Wherein P iskyReserving certain mobile reserve power for available power;
and calculating the relation between the required power of the forward flight and the forward flight speed under different tilting angles, and solving an intersection point of the available power and the required power to obtain the maximum speed boundary of the power limitation under different tilting angles.
The invention has the beneficial effects that:
(1) based on the flight mechanics model, parameters of various aspects of aircraft design are considered, including specific design parameters of rotors, wings, fuselages and the like. All design parameters can be reflected in the equilibrium equation. Therefore, compared with the existing tilting corridor calculation method of the tilting rotor aircraft, the method can reflect the actual performance of the QTR more accurately.
(2) Compared with the corridor calculated in the prior art, the tilting corridor obtained by the invention has richer connotation. The tilting corridor obtained by the prior art divides the whole flight domain into two parts according to the property of flying or not flying, and the whole flight domain is single. The tilting corridor of the invention can reflect the specific attributes of each feasible flight point, including attitude angle, required power, manipulated variable and the like, besides the basic attributes of the flight envelope.
(3) The invention is also suitable for the calculation of the tilting corridor under the tilting dual-rotor aircraft and other QTR operation modes, and the difference is the selection of the operation mode and the balancing amount. Therefore, the method has good universality.
(4) Based on the calculation method, an optimal tilting transition path/area can be found, and important reference is provided for the control strategy and flight control design. The minimum absolute value of the pitch angle is taken as an optimization target, and the machine body can tilt according to the path, so that the minimum waste resistance of the machine body can be ensured, the attitude is most stable, and the tilting is safer; and with the minimum power demand as an optimization target, finding a tilting path with the minimum power demand.
Drawings
FIG. 1 is a schematic diagram of QTR configuration;
FIG. 2 is a schematic view of a rotor model;
FIG. 3 is a schematic view of a rotor nacelle aerodynamic model;
FIG. 4 is a flow chart of the calculation of the QTR band periodic pitch-changing tilting corridor provided by the invention;
FIG. 5 is a curve of the optimized longitudinal cyclic pitch variation B1c _ opt with the change of the flying speed at different tilting angles
Figure BDA0001725320280000063
A schematic diagram;
FIG. 6 is a posture correction
Figure BDA0001725320280000061
A tilting transition region schematic diagram;
FIG. 7 is a schematic diagram of a set of tilting regions under a QTR band cyclic pitch control mode calculated by the invention;
FIG. 8 is a forward power demand fly speed curve for different tilt angles
Figure BDA0001725320280000062
A schematic diagram;
FIG. 9 is a schematic view of a rotor power limiting high speed boundary at different attitude angles and different tilt angles;
fig. 10 is a schematic diagram of the calculated results of the tilting corridor calculated by the present invention.
Detailed Description
In order to facilitate understanding of those skilled in the art, the present invention will be further described with reference to the following examples and drawings, which are not intended to limit the present invention.
Referring to fig. 1, the method for calculating the tilting corridor of the tilting quad-rotor aircraft with constant rotating speed and periodic variable pitch of the invention comprises the following steps:
1) establishing a flight mechanics model of each part of the aircraft, converting all forces and moments into a body axis system, and establishing a six-degree-of-freedom balance equation;
2) selecting a tilting operation mode: the fixed rotating speed and the periodic variable pitch control are carried out;
3) solving the upper and lower boundaries of a pitch angle by taking the lift force characteristic of the wing as a limiting factor; carrying out optimal balancing solution to obtain the manipulated variable and the attitude angle under different state conditions; finding a minimum speed limit boundary and a high speed limit boundary corresponding to the wing lift characteristic;
4) and solving the required power at different state points to obtain the available power limit high-speed boundary.
The flight mechanics model specifically comprises a rotor wing model, a nacelle model and a fuselage model. And carrying out optimized balancing calculation to obtain the balancing amount in any state.
The study object is a small QTR aircraft, and the main parameters are shown in Table 1 as follows:
TABLE 1
Figure BDA0001725320280000071
In order to research the tilting corridor corresponding to different QTR operation modes, a six-degree-of-freedom flight dynamics model needs to be established.
Referring to fig. 2, the rotor model is specifically established as follows:
a. inflow model
The induced velocity is expressed as a function of rotor spanwise position and azimuth in the form of a first order fourier series:
Figure BDA0001725320280000072
in the formula, λ0，λ1c，λ1sThe two components are respectively equal inflow, first-order longitudinal inflow component and first-order transverse inflow component of the rotor wing;
Figure BDA0001725320280000073
in the formula, a matrix M represents the influence of air inertia and represents the time delay of inflow; the V matrix is a mass flow matrix;
Figure BDA0001725320280000074
the matrix is a static relation between the disturbance induction speed and the disturbance pneumatic load; cT，CL，CMThe rotor wing lift coefficient, the pneumatic rolling moment and the pneumatic pitching moment are respectively;
considering the interference between the front rotor and the rear rotor, introducing the interference factor of the front rotor and the rear rotor, and the inflow ratio lambda of the front rotor and the rear rotorFAnd λRThe calculation formula is as follows:
Figure BDA0001725320280000075
in formula (II), lambda'F0And λ'R0The free incoming flow ratio of the front rotor wing and the rear rotor wing is respectively; v isF0V and vR0The induced speeds of the front and rear rotors are respectively; dFRFAnd dFFRInterference factors of the rear-to-front rotor and the front-to-rear rotor are respectively;
b. waving motion model
The dynamic equation of the plane of the trajectory of the rotor disk under the rotor shafting is expressed as follows:
Figure BDA0001725320280000081
in the formula, a0、a1s、b1sRespectively is a rotor flapping taper angle, a back chamfer angle and a side inclination angle;
Figure BDA0001725320280000082
is a rotor damping matrix;
Figure BDA0001725320280000083
is a rotor stiffness matrix;
Figure BDA0001725320280000084
is a rotor excitation vector;
c. rotor aerodynamic force
According to the rotor wing element momentum theory, obtaining the tension T of a rotor wing under a hub wind axis system, the backward force H of the rotor wing, the lateral force Y of the rotor wing, the pitching moment M of the rotor wing, the rolling moment L of the rotor wing and the counter torque Q of the rotor wing; converting the force and moment calculated under the rotor wind axis system to the body axis system;
the establishment of the wing model specifically comprises the following steps:
rotor wakes can have severe disturbances to the wing, dividing the wing-nacelle aerodynamic force into two parts: one part is wing aerodynamic force influenced by rotor wake, and the other part is wing aerodynamic force not influenced by rotor wake; area S of wing in slipstream partwssAnd an area S in the free stream partwfsThe calculation formulas of (A) and (B) are respectively as follows:
Figure BDA0001725320280000085
in the formula, thetanIs a tilt angle; sWIs the wing area; sssmax＝2ηssRc，ηssIs a rotor wing slipstream correction factor, R is the rotor wing radius, and c is the wing average aerodynamic chord length; 1.386, 3.114; mu is the rotor wing advancing ratio; mu.smaxThe maximum advance ratio of the rotor tail deviating out of the wing;
the wing lift and drag are respectively:
Figure BDA0001725320280000086
in the formula, qwDynamic pressure of incoming flow at the wing; cLw，CDwThe lift and drag coefficients of the wing are respectively;
the air velocity of the wing part acted by the rotor wake needs to be added with the influence of the rotor wake on dynamic pressure and lift drag coefficient; and the aerodynamic force of the wings is transmitted to the lower part of the machine body shafting.
The establishment of the rotor wing nacelle model specifically comprises the following steps:
rotor nacelle is tilting in-process, and the area of facing the wind changes, and aerodynamic resistance is the function of nacelle angle of attack, ignores other aerodynamic force and moment, as shown in fig. 3:
Dn＝4qnCDn[SnTopcos(αn)+SnSidesin(αn)] (7)
in the formula, qnThe pressure of the incoming flow at the nacelle; q is a nacelle drag coefficient; snTopThe area of the top of the nacelle; snSideIs the nacelle side area; alpha is alphanIs the nacelle angle of attack;
the establishment of the fuselage model specifically comprises the following steps:
the aerodynamic force of the fuselage is subjected to a wind tunnel test to obtain a group of airframe aerodynamic coefficients expressed in a dimensionless form, namely a resistance coefficient CDbCoefficient of lift CLbCoefficient of lateral force CSbRolling moment coefficient CMxbPitching moment coefficient CMybAnd yaw moment coefficient CMzb(ii) a The interference of a rotor wing and a wing on a fuselage is not considered;
the airframe aerodynamic force is expressed as:
Figure BDA0001725320280000091
in the formula IbFor a characteristic length of the fuselage, SbIs a characteristic area, qbThe pressure is applied to the fuselage.
Mixing and operating: QTR is the mixture of helicopter and fixed-wing aircraft, and establishes a mixed control model to solve the problem of controlRedundancy problems. By throttle/collective lever deltacolLongitudinal push rod deltalonTransverse push rod deltalatPedal deltapedThe vertical channel, the longitudinal channel, the transverse channel and the course are controlled in a mixed mode; the controlled variables are the total pitch, the longitudinal and transverse cyclic variable pitch, the four aileron deflection angles or the subsets thereof of the four ailerons, and the number of the controlled variables is set as n according to the specific operation modecThen, there are:
Figure BDA0001725320280000092
wherein u is a 4 × 1-dimensional manipulated variable vector; c is ncX 1 control variable vector; g is ncA gain matrix of x 4-dimensional manipulated variable to controlled variable;
the W weight distribution matrix is a function along with the tilting angle, and the weight changes according to a sine rule; using the controlled quantity as the total distance of four rotors
Figure BDA0001725320280000093
And four aileron deflection angles deltaflap_iA total of 8 control variables are taken as examples, i is 1,2,3, 4; then there are:
Figure BDA0001725320280000101
the weight distribution matrix is different according to different manipulation modes and is determined according to the control law.
Balance equation and solution: substituting the aerodynamic force and the moment obtained by the aerodynamic models of the QTR parts into an organism motion equation to obtain a nonlinear flight dynamics model expressed in the form of a first-order differential equation:
Figure BDA0001725320280000102
in the formula, y represents the total state amount, and t represents time.
The tilting operation mode comprises the following steps: the forward force of the front flight of the aircraft is generated by longitudinal periodic variable pitch, and is provided by tilting of the nacelle after the maximum limit of the periodic variable pitch is reached; pitch attitude can also provide a portion of the forward force; the pitching attitude of the body is mainly maintained by the differential action of the total distance between the front rotor and the rear rotor; in a helicopter mode and a large-inclination-angle tilting state, the forward flying speed caused by periodic pitch variation is obviously changed; with the gradual reduction of the tilting angle of the nacelle, the forward flight speed generated by unit period pitch change is smaller and smaller, and the control efficiency of the period pitch change is smaller and smaller until the forward flight speed change is basically not generated by the period pitch change in the fixed wing mode. The fixed speed band periodic pitch control mode is shown in table 2 as follows:
TABLE 2
Figure BDA0001725320280000103
Limiting the lift characteristics of the wings:
the lift characteristics of the wing are key factors affecting QTR flight. When the wing attack angle exceeds the wing stall attack angle alphaw_stallWhen the airplane flies, the airplane wings are considered to be stalled and can not fly normally; when the wing angle of attack is less than zero lift angle of attack alphaw0When the aircraft is in flight, the aircraft generates negative lift force, becomes a load and is very dangerous in flight state.
(a) Feasible pitching attitude angle
Pitch angle of fuselage
Figure BDA0001725320280000104
The method is very important:
the pitch angle of the fuselage is related to the attack angle of the fuselage, so that the size of the waste resistance of the fuselage is influenced;
the pitch angle of the fuselage is related to the attack angle of the wings, and the attack angle of the wings directly determines the lift-drag characteristic of the wings;
the pitch angle of the fuselage influences the attack angle of a paddle disk of the rotor wing, so that the direction of the resultant force of the rotor wing is changed, and the tension of the rotor wing is also influenced by influencing the inflow ratio of the rotor wing;
the fuselage pitch angle affects the component of the total body weight under the fuselage axis.
The pitch angle of the fuselage according to the above analysis
Figure BDA0001725320280000112
The magnitude of (b) is related to the magnitude and direction of aerodynamic force of the fuselage, wings and rotors, and is also related to the resolution of gravity under the body axis system. It involves every force within the equilibrium equation. Obviously, the pitch angle must be limited to a certain range.
Too large of a head-lowering, i.e.
Figure BDA0001725320280000113
If the wing lift force is equal to the wing lift force, a larger dynamic pressure is needed until the lift force coefficient is 0, and the dynamic pressure is larger than the wing, so that the wing cannot generate the lift force, and the maximum speed boundary is limited corresponding to the lift force characteristic of the wing. In addition, if the head-up is too large, i.e. too large
Figure BDA0001725320280000114
If the wing lift force exceeds a certain limit, the wing lift force drops suddenly due to airflow separation to generate an attack angle stall, and a minimum speed boundary is limited corresponding to the wing lift force characteristic.
Wing angle of attack exceeds wing stall angle of attack alphaw_stallWhen the airplane flies, the airplane wings are considered to be stalled and can not fly normally; when the wing angle of attack is less than zero lift angle of attack alphaw_zeroIn time, the wing cannot generate lift and becomes a load. The pitch angle range can be determined by:
Figure BDA0001725320280000111
in the formula, alphafw0，αrw0The initial stagger angles of the front and rear wings, respectively.
(b) Wing lift characteristic limit calculation process
The calculation flow is shown in fig. 4. And selecting a group of target pitch angles and tilt angles to be calculated, and respectively carrying out balancing calculation on each target pitch angle and each target tilt angle. And (4) performing optimization calculation by taking the trim pitch angle as an optimization target to obtain an optimized longitudinal periodic variable pitch, and then performing interpolation to determine the lowest speed and the highest speed in the state.
(c) Feasible tilting zone
The longitudinal cyclic pitch at different tilt angles is shown in fig. 5.
Under the helicopter mode and the large-inclination-angle tilting state, the forward flying speed caused by periodic pitch variation is obviously changed. As the tilting angle of the nacelle is gradually reduced, the forward flying speed generated by the unit-period pitch change is smaller and smaller, because the operation efficiency of the period pitch change is smaller and smaller, and the forward flying speed change is basically not generated by the period pitch change in the fixed wing mode.
As shown in fig. 6, at a pitch angle
Figure BDA0001725320280000115
For example, in the area, the QTR can be adjusted through periodic pitch change, total pitch differential and nacelle tilting, so that the pitch angle of the machine body is always 0 degrees, and the waste resistance is minimum. All the target attitudes are put together after being calculated, and as a result, as shown in fig. 7, the union of all the feasible flight domains is taken as the limit boundary of the lift characteristic of the wing.
Power limited high speed boundary: the aircraft must meet the power requirements in order to reach a certain state;
during the trim calculation, the torque of the rotor is extracted and checked to be within the available power range. The power required by the whole machine is as follows:
Figure BDA0001725320280000121
in the formula, P is required power; qiThe reaction torque of the ith auxiliary rotor wing;
then only the check is needed: p is less than or equal to Pky(ii) a Wherein P iskyReserving certain mobile reserve power for available power;
the relationship between the required power of the forward flight and the forward flight speed at different tilting angles is calculated, and is shown in fig. 8.
And (4) solving an intersection point of the available power and the required power to obtain the maximum speed boundary of the power limit under different tilting angles. As shown in table 3, the maximum velocity boundary values for different pitch angles and different tilt angles are as follows:
TABLE 3
Figure BDA0001725320280000122
The power envelopes at the different attitude angles are shown in fig. 9. The power limiting tilt envelope is the union of the lower bounds of all pitch angles. From the above conclusion, the boundary of the 0 ° pitch angle includes the rest of the boundaries, so the boundary line corresponding to the 0 ° pitch angle is the final power-limiting speed envelope.
The four boundaries calculated above are plotted on a graph to obtain a complete tilting corridor, as shown in fig. 10. The tilt region for attitude optimization is also given: the attitude angle is 0, and the waste resistance is minimum. In consideration of flight safety, the maximum speed corresponding to a tilt angle of 45 degrees is taken as the suspension speed. The aircraft cannot exceed the stopping speed during the tilting process.
While the invention has been described in terms of its preferred embodiments, it will be understood by those skilled in the art that various changes in form and details may be made therein without departing from the spirit and scope of the invention.

Claims (5)

1. The utility model provides a four rotor crafts that vert corridor calculation method that calms rotational speed area cycle displacement, its characterized in that includes the step as follows:
1) establishing a flight mechanics model of each part of the aircraft, converting all forces and moments into a body axis system, and establishing a six-degree-of-freedom balance equation;
2) selecting a tilting operation mode;
3) solving the upper and lower boundaries of a pitch angle by taking the lift force characteristic of the wing as a limiting factor; carrying out optimal balancing solution to obtain the manipulated variable and the attitude angle under different state conditions; finding a minimum speed limit boundary and a high speed limit boundary corresponding to the wing lift characteristic;
4) obtaining an available power limit high-speed boundary by solving the required power under different state points;
the flight mechanics model in the step 1) comprises: the model comprises a rotor wing model, a rotor wing nacelle model and a fuselage model;
the step 3) is specifically as follows: wing angle of attack exceeds wing stall angle of attack alphaw_stallWhen the airplane flies, the airplane wings are considered to be stalled and can not fly normally; when the wing angle of attack is less than zero lift angle of attack alphaw_zeroWhen the wing is in use, the wing can not generate lift force and becomes a load; the pitch angle θ range is determined by:
Figure FDA0002926863070000011
in the formula, alphafw0，αrw0Respectively the initial mounting angles of the front and rear wings;
selecting a group of target pitch angles and tilt angles to be calculated, and respectively carrying out balancing calculation on each target pitch angle and each target tilt angle; carrying out optimization calculation by taking the trim pitch angle as an optimization target to obtain an optimized longitudinal periodic variable pitch; then, determining the lowest speed and the highest speed in the state through interpolation;
putting all the target attitudes together after completing calculation, and taking a union of all feasible flight areas as a limit boundary of the lift characteristic of the wing;
the step 4) is specifically as follows: in the process of trim calculation, the torque of the rotor wing is extracted to calculate the required power, and the required power of the whole machine is as follows:
Figure FDA0002926863070000012
in the formula, P is required power; qiThe reaction torque of the ith auxiliary rotor wing;
then only the check is needed: p is less than or equal to Pky(ii) a Wherein P iskyReserve a certain reserve of power for usePower;
and calculating the relation between the required power of the forward flight and the forward flight speed under different tilting angles, and solving an intersection point of the available power and the required power to obtain the maximum speed boundary of the power limitation under different tilting angles.
2. The method for calculating the pitch corridor of a fixed-speed and periodically variable tilting quad-rotor aircraft according to claim 1, wherein the step 1) further comprises:
11) the rotor model building method comprises the following steps:
a. inflow model
Induction speed
Figure FDA0002926863070000021
Expressed as rotor spanwise position in the form of a first order Fourier series
Figure FDA0002926863070000022
And azimuth ψ:
Figure FDA0002926863070000023
in the formula, λ0，λ1c，λ1sThe two components are respectively equal inflow, first-order longitudinal inflow component and first-order transverse inflow component of the rotor wing;
Figure FDA0002926863070000024
in the formula, the matrix M is an inertia matrix and reflects the time delay of inflow; the V matrix is a mass flow matrix;
Figure FDA0002926863070000025
the matrix is a static relation matrix between the disturbance induction speed and the disturbance pneumatic load; cT，CL，CMThe rotor wing lift coefficient, the pneumatic rolling moment and the pneumatic pitching moment are respectively;
considering the interference between the front rotor and the rear rotor, introducing the interference factor of the front rotor and the rear rotor, and the inflow ratio lambda of the front rotor and the rear rotorFAnd λRThe calculation formula is as follows:
Figure FDA0002926863070000026
in formula (II), lambda'F0And λ'R0The free incoming flow ratio of the front rotor wing and the rear rotor wing is respectively; v isF0V and vR0The induced speeds of the front and rear rotors are respectively; dFRFAnd dFFRInterference factors of the rear-to-front rotor and the front-to-rear rotor are respectively;
b. waving motion model
The dynamic equation of the plane of the trajectory of the rotor disk under the rotor shafting is expressed as follows:
Figure FDA0002926863070000027
in the formula, a0、a1s、b1sRespectively is a rotor flapping taper angle, a back chamfer angle and a side inclination angle;
Figure FDA0002926863070000028
is a rotor damping matrix;
Figure FDA0002926863070000029
is a rotor stiffness matrix;
Figure FDA00029268630700000210
is a rotor excitation vector;
c. rotor aerodynamic force
According to the rotor wing element momentum theory, obtaining the tension T of a rotor wing under a hub wind axis system, the backward force H of the rotor wing, the lateral force Y of the rotor wing, the pitching moment M of the rotor wing, the rolling moment L of the rotor wing and the counter torque Q of the rotor wing; converting the force and moment calculated under the rotor wind axis system to the body axis system;
12) the establishment of the wing model specifically comprises the following steps:
the aerodynamic force of the wing-nacelle is divided into two parts: one part is wing aerodynamic force influenced by rotor wake, and the other part is wing aerodynamic force not influenced by rotor wake; area S of wing in slipstream partwssAnd an area S in the free stream partwfsThe calculation formulas of (A) and (B) are respectively as follows:
Figure FDA0002926863070000031
in the formula, thetanIs a tilt angle; sWIs the wing area; sssmax＝2ηssRc，ηssIs a rotor wing slipstream correction factor, R is the rotor wing radius, and c is the wing average aerodynamic chord length; 1.386, 3.114; mu is the rotor wing advancing ratio; mu.smaxThe maximum advance ratio of the rotor tail deviating out of the wing;
the wing lift and drag are respectively:
Figure FDA0002926863070000032
in the formula, qwDynamic pressure of incoming flow at the wing; cLw，CDwThe lift and drag coefficients of the wing are respectively;
13) the establishment of the rotor wing nacelle model specifically comprises the following steps:
rotor nacelle is changeing the in-process, and the area that faces the wind changes, and aerodynamic resistance is the function of nacelle angle of attack, ignores other aerodynamic force and moment:
Dn＝4qnCDn[SnTopcos(αn)+SnSidesin(αn)] (7)
in the formula, qnThe pressure of the incoming flow at the nacelle; cDnIs the nacelle drag coefficient; snTopThe area of the top of the nacelle; snSideIs the nacelle side area; alpha is alphanIs the nacelle angle of attack;
14) the establishment of the fuselage model specifically comprises the following steps:
obtaining a group of airframe aerodynamic coefficients expressed in a dimensionless form, namely resistance coefficients C through a wind tunnel testDbCoefficient of lift CLbCoefficient of lateral force CSbRolling moment coefficient CMxbPitching moment coefficient CMybAnd yaw moment coefficient CMzb(ii) a The interference of a rotor wing and a wing on a fuselage is not considered;
the airframe aerodynamic force is expressed as:
Figure FDA0002926863070000033
in the formula IbIs a characteristic length of the fuselage, AbIs a characteristic area, qbThe pressure is applied to the fuselage.
3. The method for calculating the tilting corridor of the fixed-speed and periodically-variable tilting quad-rotor aircraft according to claim 1, wherein the step 1) specifically comprises the following steps: establishing a hybrid control model through an accelerator/collective pitch rod deltacolLongitudinal push rod deltalonTransverse push rod deltalatPedal deltapedThe vertical channel, the longitudinal channel, the transverse channel and the course are controlled in a mixed mode; the controlled variables are the total pitch, the longitudinal and transverse cyclic variable pitch, the four aileron deflection angles or the subsets thereof of the four ailerons, and the number of the controlled variables is set as n according to the specific operation modecThen, there are:
Figure FDA0002926863070000041
wherein u is a 4 × 1-dimensional manipulated variable vector; c is ncX 1 control variable vector; g is ncA gain matrix of x 4-dimensional manipulated variable to controlled variable;
the W weight distribution matrix is a function along with the tilting angle, and the weight changes according to a sine rule; using the controlled quantity as the total distance of four rotors
Figure FDA0002926863070000044
And four aileron deflection angles deltaflap_iA total of 8 control variables are taken as examples, i is 1,2,3, 4; then there are:
Figure FDA0002926863070000042
4. the method for calculating the tilting corridor of the fixed-speed and periodically-variable tilting quad-rotor aircraft according to claim 1, wherein the step 1) specifically comprises the following steps: substituting the aerodynamic force and the moment obtained by the aerodynamic models of all parts of the four-rotor aircraft into an organism motion equation to obtain a nonlinear flight dynamics model expressed in the form of a first-order differential equation:
Figure FDA0002926863070000043
where y represents the total state quantity, u is a 4 × 1-dimensional manipulated variable vector, and t is time.
5. The method for calculating the tilting corridor of the fixed-speed and periodically-variable tilting quad-rotor aircraft according to claim 1, wherein the tilting operation mode selected in the step 2) is specifically a fixed-speed and periodically-variable tilting operation mode: the forward force of the front flight of the aircraft is generated by longitudinal periodic variable pitch, and is provided by tilting of the nacelle after the maximum limit of the periodic variable pitch is reached; gradually adding ailerons for operation; pitch attitude can also provide a portion of the forward force; the pitching attitude of the airframe is mainly maintained by the collective pitch differential of the front and rear rotors.
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Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101643116A (en) * 2009-08-03 2010-02-10 北京航空航天大学 Tiltrotor controlled by double-propeller vertical duct
CN103144781A (en) * 2012-12-21 2013-06-12 南京航空航天大学 Method for determining transient process switching corridor of tilt rotor unmanned aircraft
CN103803079A (en) * 2012-11-12 2014-05-21 上海市闵行区知识产权保护协会 Small-size rotor aircraft
CN106828885A (en) * 2016-12-30 2017-06-13 上海牧羽航空科技有限公司 A kind of use jet form control driftage and the tiltrotor of pitching

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2982964B1 (en) * 2011-11-23 2013-11-22 Eurocopter France METHOD FOR ASSISTED STEERING OF A ROTARY TURNING AIRCRAFT COMPRISING AT LEAST ONE PROPELLANT PROPELLER, ASSISTED STEERING DEVICE AND AIRCRAFT

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101643116A (en) * 2009-08-03 2010-02-10 北京航空航天大学 Tiltrotor controlled by double-propeller vertical duct
CN103803079A (en) * 2012-11-12 2014-05-21 上海市闵行区知识产权保护协会 Small-size rotor aircraft
CN103144781A (en) * 2012-12-21 2013-06-12 南京航空航天大学 Method for determining transient process switching corridor of tilt rotor unmanned aircraft
CN106828885A (en) * 2016-12-30 2017-06-13 上海牧羽航空科技有限公司 A kind of use jet form control driftage and the tiltrotor of pitching

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