CN113868754B - Combined helicopter control distribution and optimal transition route design method - Google Patents
Combined helicopter control distribution and optimal transition route design method Download PDFInfo
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Abstract
The invention provides a method for controlling and distributing a combined helicopter and designing an optimal transitional route, which establishes a flight mechanics model of HCH (hydrogen chloride), and specifically comprises a rotor wing model, a nacelle model and a fuselage model. And performing optimization balancing calculation to obtain balancing quantity in any state. The control strategy of the composite high-speed helicopter at each speed is provided, and the control distribution coefficient among different control mechanisms and the transition route of the pitching attitude angle in the transition process are obtained through design and optimization according to the control strategy, so that each control amount is smoothly transited and the attitude amount is maintained within a reasonable range. The method is based on the flight mechanics model, and the calculation method can be fed back to the overall design link to provide reference for overall parameter optimization of the aircraft. The transition route and the distribution coefficient obtained by the invention are subjected to balancing and are compared with the power and the control amount obtained by balancing the fixed pitch angle, so that the method has the advantage of smaller power, and the control amount can be smoothly and continuously transited, thereby being more in line with the control habit of a driver.
Description
Technical Field
The invention relates to the field of flight mechanics and control of a compound rotor aircraft, in particular to a method for controlling distribution and designing an optimal transitional route of a compound helicopter.
Background
The composite high-speed helicopter propelled by double propellers is a design concept for realizing high-speed flight of the helicopter, the configuration is shown in figure 1, the aircraft is provided with two wings on two sides of a fuselage on the basis of a conventional helicopter, a tail rotor is omitted, a tail rotor is additionally arranged, and a pair of propulsion propellers are arranged on two sides of the wing. The addition of a pair of propulsion propellers, ailerons, elevators and rudders means that four operating variables are added, and the resulting problem of operating redundancy presents a difficult problem for the design and trim analysis of a compound helicopter control system
In order to delay the problems of airflow compressibility of the forward blade and stall of the backward blade when the conventional helicopter flies forward, the HCH introduces thrust composite to enable the wing to be used as a main rotor wing for lifting and unloading, and reduces the rotating speed so as to delay the inherent problems of the conventional helicopter. However, the compound thrust brings about the problem of strong coupling in operation, and the addition of a large number of operating mechanisms brings about the redundant control problem of the operating amount, the operating mechanisms of the helicopter and the fixed wing exist on the same attitude control loop, and the method for distributing the control surface of the wing and the control surface of the helicopter is provided with very important research value.
Modeling and stability characteristic analysis are carried out on the composite high-speed helicopter at home and abroad, the influence of parameters of the flyer on performance in the design process is analyzed, and the double-propeller propulsion type composite helicopter has the public work in the aspects of transition process flight control design and transition route. The method comprises the steps that a literature establishes a flight mechanics mathematical model of an X2 and X3 composite high-speed helicopter structure by adopting a given flight mechanics model of a conventional helicopter and introducing the combination of a propeller and a wing into a design, compares the model with the conventional helicopter structure, researches balancing characteristics, stability and operability of the composite helicopter, and discusses the control quality problem of the composite helicopter on the basis of the balancing characteristics, stability and operability; literature studies the impact of overall parameters on the performance of a compound helicopter; the literature takes the lift force distribution amount of the composite helicopter as an optimization variable, and takes the lowest required power as an optimization target for optimization. To obtain a law of lift distribution at the lowest required power. The composite helicopter is still under exploration at home and abroad, and the distribution problem of the two mechanisms is rarely reported. Compared with a conventional helicopter, the HCH is added with a pair of propelling propellers, a pair of wings and a pair of tail wings, and the pair of tail wings are arranged below the tail of the rotor wing, so that force changes in the flight process of the HCH cannot be accurately calculated by using a helicopter model alone. A flight mechanics model of HCH needs to be established, and interference of each part is considered; the new manipulation quantity is increased, and the 6 dynamic relations cannot solve 7 unknown quantities of HCH trimming solution, so that a pitch angle is set as a fixed value to carry out trimming calculation, and a pitch angle transition route design problem at different speeds is generated. The HCH low-speed stage mainly uses the main rotor wing to provide lifting force to overcome gravity, at this moment, the control efficiency of the control surface of the wing is very small, the rudder efficiency of the wing is gradually increased along with the increase of speed, and for this relationship, a distribution strategy needs to be designed to realize the distribution between two mechanisms.
Based on the background, the HCH manipulation allocation and optimal transition route design method has important significance on overall design, component design, flight quality, manipulation strategy and control system design.
Disclosure of Invention
The invention provides a method for designing a combined helicopter control distribution and optimal transition route, which aims to solve the problems in the prior art, and design parameters can be reflected in a balance equation based on a flight mechanics model to provide a reference for the optimization of overall parameters of an aircraft. The transition route and the distribution coefficient obtained by the invention are subjected to balancing and are compared with the power and the control amount obtained by balancing the fixed pitch angle, so that the method has the advantage of smaller power, and the control amount can be smoothly and continuously transited, thereby being more in line with the control habit of a driver.
The invention comprises the following steps:
1) Establishing a flight mechanics model of each component, converting all forces and moments into a body shafting, and establishing a six-degree-of-freedom balance equation;
2) Determining maneuvering strategies under different flying speeds, and determining maneuvering mechanisms participating in maneuvering under different flying speeds, wherein the maneuvering strategies under different flying speeds comprise a hovering and low-speed maneuvering strategy, a high-speed forward flying maneuvering strategy and a transitional forward flying maneuvering strategy, the maneuvering mechanisms under the hovering and low-speed maneuvering strategies comprise a main rotor wing and a double-propulsion propeller, the flying speeds are 0m/s to 10m/s, the main rotor wing is subjected to collective maneuvering, longitudinal periodic variable-pitch maneuvering and transverse periodic variable-pitch maneuvering, and the double-propulsion propeller comprises an average pitch and a differential pitch; the control mechanism of the high-speed forward flight control strategy comprises a double-propulsion propeller, fixed wings and a tail wing, wherein the flight speed interval is more than 45m/s, the fixed wings are an elevator, an aileron, a rudder, a pitch channel and a pitching channel, and the double-propulsion propeller comprises an average pitch; the control mechanism of the transition front flight control strategy comprises a double-propulsion propeller, a main rotor wing, a fixed wing and a tail wing, wherein the speed interval of the transition stage is 10m/s-45m/s, the fixed wing is an elevator, an aileron, a rudder, a pitch channel and a pitching channel, the main rotor wing is subjected to collective pitch control, longitudinal periodic pitch control and transverse periodic pitch control, and the double-propulsion propeller comprises an average pitch and a differential pitch;
3) Designing a redundant control distribution coefficient taking the speed as a reference quantity, judging whether the pitch angle meets boundary conditions, and optimizing to obtain the pitch angle in each state to obtain a transition route of the compound helicopter;
4) Determining an optimization strategy of different flight speed stages according to different maneuvering strategies:
4.1 An optimization strategy of the transitional fly-forward manipulation strategy;
the pitch angle is optimized for the first time by using a Newton method, the distribution coefficient of the optimal maneuvering efficiency is used as the initial value of the distribution coefficient, one distribution coefficient corresponding to each flying speed is used for maximizing the current maneuvering efficiency, and the maneuvering coefficient is as follows:
wherein the method comprises the steps of
Expressed as pitching moment caused by the unit steering amount, < >>
Respectively a unit longitudinal period variable pitch operation and a unit elevator operation, K' cyc 、K′ ail Expressed as a longitudinal period change distance and elevator distribution coefficient in the current state, and K 'is regulated' cyc +K′ ail ＝1；
Longitudinal period change is delta lon_cyc ＝K′ cyc *δ lon The elevator steering amount is delta lon_ail ＝K′ ail *δ lon ，δ lon Is pitch channel manipulation amount;
when the distribution coefficient is optimized to zero, the control quantity is continuously reduced along with the increase of the flying speed, and the control quantity cannot meet the boundary condition through optimizing the distribution coefficient, so that the pitch angle is optimized, and the optimal power is achieved on the premise that the pitch angle meets the boundary condition;
4.2 An optimization strategy for hover and low speed maneuver strategies;
because the flying speed is low, the air flow speed acting on the wing is too low, the aerodynamic force of each control surface is small, and the efficacy of the control surface is very low, so that the distribution coefficient between the rotor wing and the wing control surface in the low-speed stage is fixed to be 1;
in the speed range of 0-10m/s, keeping the pitching attitude to be zero, avoiding the negative pitch, and the pitch angle constraint condition of the low-speed stage is as follows:
AVER(θ)-DIFF(θ)＞0(rad)
wherein AVER and DIFF are respectively the average pitch and the differential pitch in the trim state;
4.3 An optimization strategy of the high-speed fly-forward manipulation strategy;
the wing bears 90% of lift force in the high-speed stage, and the control surface has enough control efficacy to perform attitude operation, so that the distribution coefficient between the rotor wing in the high-speed stage and the wing control surface is a fixed value;
the high-speed stage ensures that the total distance is positive, and the constraint condition of the total angle of the high-speed stage is as follows:
Col＞0(rad)
where Col is the total angle in trim state, respectively;
5) Obtaining control distribution coefficients among different control mechanisms and a pitching attitude angle transition route in the transition process;
6) The calculation method is applied to the composite helicopter aircraft for calculation.
The invention has the beneficial effects that:
1. based on the flight mechanics model, parameters of various aspects of the aircraft design are considered, including specific design parameters of the rotor, wing, fuselage, and the like. All design parameters can be reflected in the equilibrium equation. Therefore, the calculation method can be fed back to the overall design link, and a reference is provided for the overall parameter optimization of the aircraft.
2. The transition route and the distribution coefficient obtained by the invention are subjected to balancing and are compared with the power and the control amount obtained by balancing the fixed pitch angle, so that the method has the advantage of smaller power, and the control amount can be smoothly and continuously transited, thereby being more in line with the control habit of a driver.
Drawings
Fig. 1 is a compound helicopter coordinate system.
FIG. 2 is a pitch angle and pitch distribution coefficient optimization algorithm at different speeds.
Fig. 3 shows the tension change around zero pitch.
Fig. 4 is a pitch channel distribution coefficient.
Fig. 5 is a rolling channel allocation coefficient.
FIG. 6 is a heading channel allocation coefficient.
Fig. 7 is a pitch attitude angle transition route.
FIG. 8 is a graph of rotor and wing lift as a function of speed.
Fig. 9 is a longitudinal steering amount.
Fig. 10 is a lateral steering amount.
Fig. 11 is a power calculation diagram.
Fig. 12 is a propulsive propeller pitch.
Fig. 13 is a power comparison of designed versus un-designed routes.
Fig. 14 shows the redundant steering and the unassigned longitudinal steering amounts.
Detailed Description
The invention is further described below with reference to the drawings and the detailed description.
One embodiment of the invention is as follows:
1. establishing a flight dynamics model
1.1 flight dynamics modeling
The calculation of the control strategy design and the control mechanism distribution coefficient is based on a double-propeller compound helicopter verification prototype. The double-propeller composite helicopter verification prototype uses a design scheme of combining thrust and lift force, wherein the lift force is provided by a main rotor wing and a box-shaped wing connected with a fuselage, the wing forms a closed triangular structure, four wings are respectively provided with an aileron, and two ends of two pairs of wings are respectively provided with a propulsion propeller. The tail of the machine body is additionally provided with a horizontal tail wing and a vertical tail wing. The overall parameters of the composite helicopter validator used herein are shown in table 1.
Table 1 sample helicopter parameters overall
The total weight of the prototype fuselage was calculated to be 15kg, the main rotor diameter 1340mm and the propulsive pitch radius 155mm. The vertical position of the main rotor wing relative to the gravity center is 0.23m, the transverse and longitudinal positions are coincident with the gravity center, and the propulsion paddle is positioned above the gravity center by 0.08m. A composite helicopter model represented by the general parameters herein is shown in fig. 1, and an inertial coordinate system and a body coordinate system are established as shown.
The aerodynamic force of the aerodynamic components is calculated respectively to obtain aerodynamic force and moment of each component.
(1) Rotor aerodynamic force
Compared with a conventional helicopter, the modeling of the main rotor wing of the composite helicopter is the same as that of the conventional helicopter, the rotor wing model is modeled by adopting a phyllotoxin theory, the uniform inflow induction speed is solved in an iterative manner, the first-order rigidity waving of the blade is only considered, and the aerodynamic force of the micro-segment of the blade is calculated according to the phyllotoxin theory.
Wherein the method comprises the steps of
Is the tangential and normal airflow velocity of the regularized blade profile, c is the blade chord length. In addition to aerodynamic moments caused by aerodynamic forces, on blade segmentsThe moment also includes centrifugal moment, blade moment of inertia, flapwise spring moment and coriolis moment. And integrating to obtain the force and moment generated by the main rotor wing.
(2) Aerodynamic force of propulsion propeller
The modeling of the propulsion propeller is to divide the propeller blade into a section of micro-segment, the key is to calculate the speed at the micro-segment of the propeller blade, and the speed at the 1/4 chord line position under the propeller blade shafting is as follows:
v tan 、v n Is the tangential velocity and the vertical velocity of the blade micro-segment, theta twist Is the negative torque of the blade. And obtaining the lift resistance coefficient and the corresponding force and moment by a table look-up mode, and then integrating to obtain the force and moment of the whole propeller to the center of the hub.
(3) Wing, vertical tail and horizontal tail aerodynamic force
In a hovering state, the rotor wake induces aerodynamic loads to the wing, so that low-speed performance is reduced, and the speed of the wing micro-section under the body shafting is as follows:
the wing aerodynamic force has the following structure:
the current wing model uses the product of the hypothetical empirical coefficient and the main rotor wake uniform induction speed to perturb the wing, matching the hover load empirical coefficient k measured by Felker and Light in the experimental test of V-22 wing loads, selected to be 1.5, where
Is the aircraft speed, ρ is the air density, +.>
Is the chord length of the wing, C L 、C D 、C M Respectively the lift force, the resistance and the pitching moment coefficient.
And obtaining the force and moment generated by each pneumatic component of the compound helicopter, regarding the compound helicopter as a rigid body, and establishing a six-degree-of-freedom motion equation.
Wherein I is n For the inertial matrix of the aircraft, u, v, w are the velocities under the body axis, p, q, r are the angular velocities under the body axis,
is the resultant force of all pneumatic components under the body axis, and +.>
The combined moment of each aerodynamic component under the body shafting is that theta, phi and phi are Euler angles of the aircraft, and the combined helicopter steering quantity comprises a total pitch lever steering quantity (rotor collective pitch), a transverse lever steering quantity (rotor transverse cyclic pitch and aileron), a longitudinal lever steering quantity (rotor longitudinal cyclic pitch and elevator), a heading lever steering quantity (propeller differential pitch and rudder) and an average pitch lever (propeller average pitch). Wherein different operating mechanisms of the same operating quantity are combined through the distribution coefficient. When trimming calculation is carried out, the resultant force of the machine body is zero, and the combined moment of the machine body is zero, so that a motion equation set of the machine body is established as follows:
∑F x -Gsinθ＝0
∑F y +Gcosθsinφ＝0
∑F x +Gcosθcosφ＝0
∑M x ＝0
∑M y ＝0
∑M z ＝0 (8)
the number of the motion equations of the machine body is 6, the total distance rod operation amount, the transverse rod operation amount, the longitudinal operation amount, the heading rod operation amount, the average pitch operation amount, the pitch angle and the roll angle are 7, an additional state is required to be specified, and the dimension of the equation set is reduced. Because the pitch attitude angle directly influences the thrust required to be generated by the double-propulsion propeller, the lift force generated by the wings at two sides is also influenced by the attack angle, so that the pitch attitude angle is regulated in the trimming process.
Finally, the amounts to be trimmed are the collective stick manipulation amount, the lateral stick manipulation amount, the longitudinal manipulation amount, the heading stick manipulation amount, the average pitch manipulation amount, and the roll angle.
1.2 composite helicopter steering strategy
1.2.1 hover/Low speed maneuver strategy
In hover/low speed mode, the compound helicopter is similar to a conventional helicopter, with the main steering mechanisms being rotors (collective, longitudinal cyclic, transverse cyclic) and double propulsive propellers (average pitch, differential pitch). The flying speed of the mode is about 0m/s to 10 m/s. Because the compound helicopter is provided with the wing, in order to ensure that the wing can be kept in the attack angle range capable of generating lift force under different speeds, the thrust of the compound helicopter is mainly provided by double propellers, and the longitudinal period pitch variation mainly controls the pitching attitude of the compound helicopter. The composite helicopter adopting the control strategy can change the attack angle of the wing at different flying speeds, so that the wing can keep an optimal working state. The specific manipulation strategy is shown in table 2.
TABLE 2 steering strategy in Low speed/hover mode
1.2.2 high speed fly-forward maneuver strategy
When flying forward at high speed, the wing generates most of lift force, the main control mechanism is a double screw propeller, a wing and a tail wing, the flying speed interval of the mode is more than 45m/s, and the rotor wing can delay shock wave resistance at the forward blade by reducing the rotating speed of the main rotor wing. The specific manipulation strategy is shown in table 3.
TABLE 3 strategy for handling in high speed front fly mode
1.2.43 transitional fly-forward steering strategy
The rudder efficiency of the wing is gradually increased when the composite helicopter flies forward, the flight characteristics of the fixed wing are shown when the composite helicopter flies at a high speed, and the flight control strategy is required to be suitable for the flight characteristics and the control habit of the fixed wing when the composite helicopter flies forward. The main rotor wing of the compound helicopter in the transition stage also participates in the attitude control and the altitude control of the helicopter body. The speed interval in the transition stage is 10m/s-45m/s, and the wing of 10m/s starts to bear 10% of the lift force.
In the transition stage, the number of the combined helicopter steering variables is 8, and classification is shown in table 4 according to different control channels.
TABLE 4 steering strategy in transitional forward flight mode
2. Transition route and distribution coefficient optimization design
The control mode of the double-propeller propulsion compound helicopter in different stages is analyzed, but in the transition stage, the pitching channel has elevators and longitudinal period variable pitch due to redundancy of control surfaces, and the control modes are mainly converted into the elevators when the helicopter enters a fixed wing mode. Additional provision is required for the pitch angle in the trim calculation, which results in a power impact due to the greater impact of the pitch angle on the average pitch. Therefore, in order to solve the problem of operation redundancy of the compound helicopter when the speed is lower than the limit speed, the section takes a pitching channel as an example, designs a redundancy operation distribution coefficient taking the speed as a reference quantity, and optimizes and obtains the pitching angle under each state to obtain a transition route of the compound helicopter. 2.1 optimization goals and boundary conditions
The specified pitch angle transition route can influence the pitch of the propeller in the current state and the longitudinal period variable pitch, and further influence the full-machine power in the current state, so that the power is used as an optimization target to obtain the optimal pitch angle in each state.
In the control law design, it is expected that the longitudinal rod amount of the pitching channel can smoothly transition with the increase of the flying speed along with the increase of the flying speed, so that the optimization strategy is as follows: on the premise of smooth transition of the manipulation amount, the pitch angle is optimized to enable the power to be the lowest, and the distribution coefficient with the maximum manipulation efficiency at the moment is obtained.
Definition of the definition
K′ cyc 、K′ ail Expressed as a longitudinal period variable pitch and elevator distribution coefficient in the current state, the optimized objective function is to optimize the full-aircraft power of the compound helicopter, and is described as follows:
M ob ＝P(a,θ) (9)
the analysis results in that the determination of the boundary conditions in the control strategy is based on the distribution coefficient under the optimal efficacy, and the distribution coefficient is optimized, so that the rod amount of the attitude passage can be smoothly transited along with the increase of the flying speed, and the values of the boundary conditions are selected by the trimming characteristics of the helicopter, so that the rod amount of the attitude passage can be transited along the same direction along with the increase of the flying speed. The boundary conditions determined are:
The boundary conditions of the roll channel are determined as:
As the forward flying speed of the compound helicopter is increased, the total distance of the rotor is reduced, the reactive torque of the rotor is reduced, and the heading steering amount is gradually reduced to be close to 0, the heading distribution coefficient is not optimized, and is set to be the strategy of ensuring the maximum steering efficiency in order to ensure the range of the average pitch, and is kept to be 0.5 when flying forward at high speed.
2.2 optimization strategies
2.2.1 transition phase optimization strategy
The optimization algorithm optimizes the transition route with optimal power at different speeds by adopting a Newton method, takes the pitch angle as a prescribed quantity in the optimization calculation of the distribution coefficient, and the optimization of the distribution coefficient also influences the pitch angle with optimal power, so that the transition route and the distribution coefficient which meet the boundary condition and enable the objective function to reach the optimal are obtained through repeated cyclic optimization. From the above analysis, fig. 2 shows an algorithm flow chart.
The first optimization of pitch angle using newton's method requires the use of the distribution coefficient of optimal steering efficiency as the distribution coefficient initial value. One distribution coefficient for each flight speed is to maximize the current maneuvering efficiency. The steering coefficient is:
wherein the method comprises the steps of
Expressed as pitching moment caused by the unit steering amount, < >>
Respectively a unit longitudinal period variable pitch operation and a unit elevator operation, K' cyc 、K′ ail Expressed as a longitudinal period change distance and elevator distribution coefficient in the current state, and K 'is regulated' cyc +K′ ail ＝1。
Longitudinal period change is delta lon_cyc ＝K′ cyc *δ lon The elevator steering amount is delta lon_ail ＝K′ ail *δ lon ，δ lon Is the pitch channel handling.
When the distribution coefficient is optimized to zero, the control quantity is continuously reduced along with the increase of the flying speed, and the control quantity cannot meet the boundary condition by optimizing the distribution coefficient, the pitch angle is optimized, and the pitch angle reaches the optimal power on the premise of meeting the boundary condition.
2.2.2 Low speed stage optimization strategy
Because the flying speed is low, the air flow speed acting on the wing is too low, the aerodynamic force of each control surface is small, and the efficacy of the control surface is very low, so that the distribution coefficient between the rotor wing and the control surface of the wing in the low-speed stage is fixed to be 1.
In the speed range of 0-10m/s, keeping the pitching attitude to be zero, and needing the left-side propeller to provide negative thrust to balance the reactive torque of the main rotor, wherein the forward flying speed and the induced speed of the left-side propeller are opposite in direction, and when the current flying speed reaches the induced speed of 0.7 times, the left-side propeller can enter a vortex ring state, so that the pull loss exists in the propeller, and the flying safety in transition is influenced; FIG. 3 shows the situation that the screw pitch changes from negative to positive and the screw generates a pulling force in the forward incoming flow of 8m/s, and the drawing shows that the screw induction speed is reverse and the pulling force changes severely near the zero screw pitch.
In consideration of aerodynamic problems caused by negative pitches, the negative pitches should be avoided as much as possible. The pitch angle constraints for the low speed phase are:
AVER(θ)-DIFF(θ)＞0(rad) (13)
wherein AVER and DIFF are the average pitch and the differential pitch in the trim state, respectively.
2.2.3 high speed stage optimization strategy
The wing bears 90% of lift force in the high-speed stage, and the control surface has enough control efficacy to perform attitude operation, so that the distribution coefficient between the rotor wing in the high-speed stage and the control surface of the wing is a fixed value.
In the high-speed stage, wing lift force is increased, a main rotor wing even presents a negative pitch state, the fact that the rotor wing blows and swings to be related to the taper angle of the rotor wing is noticed, the rotor wing can be changed from backward tilting to forward tilting, the change of forward force of the rotor wing can be influenced, and the total distance is ensured to be positive as much as possible in consideration of aerodynamic force caused by the change of the taper angle. The total pitch angle constraint for the high speed phase is:
Col＞0(rad) (14)
where Col is the total pitch angle in trim state, respectively.
3. Optimizing results and discussion
3.1 distribution coefficient and pitch transition route results
According to the design of the control strategy, the distribution coefficient design of the pitching channel and the rolling channel is carried out, and the distribution coefficient with continuous input quantity of the attitude angle ring and optimal power and the transition route of the compound helicopter are obtained.
The distribution coefficients of the longitudinal and rolling channels and the pitching attitude angles are shown in figures 4-7, and the distribution coefficient of the heading channel adopts a linear decreasing strategy. The pitch angle transition route shows a rule of reducing the approach level along with the increase of the flying speed; at low speeds, rotor cyclic variation takes a main role, and as the speed increases, rotor cyclic variation distribution coefficients gradually decrease, and fixed wing control surface distribution coefficients gradually increase.
3.2 numerical calculation
According to the solved pitching attitude angle, during balancing calculation, the pitching angle is fixed according to the transition route, so that the amount to be balanced is reduced to 6, and balancing calculation is completed.
Fig. 8 shows a graph of the variation of the lift assumed by the rotor and the wing with increasing speed, the rotor assuming the main lift at low speed, the lift of the wing increasing with increasing speed, unloading the lift for the rotor. The rotor bears the reduced lift and the resultant distance value is reduced.
The steering amount and the composite helicopter power of each channel are obtained through solving and are shown in figures 9-12. The optimized distribution coefficient and the pitching attitude angle can enable the control quantity of each attitude channel to smoothly transition along with the flying speed, and the power optimization is realized. In which the longitudinal steering amount obtained in fig. 9 gradually transitions from longitudinal cyclic to elevator as the speed increases, and the longitudinal steering amount can smoothly transition during acceleration. The lateral steering amount obtained in fig. 10 also achieves a smooth transition and a reasonable distribution of the two steering mechanisms.
In addition, it can be derived from fig. 12 that as the forward speed increases, the pitch of the propellers on the left and right sides increases, with the inflow speed of the propellers being greater, the pitch being greater while flying at high speeds, but the effective angle of attack being maintained within a reasonable range, with 50m/s remaining at an effective angle of attack of 10 °, and the stall condition not being reached. The pitch is positive in the low-speed section, and the propeller does not encounter the vortex ring state in the acceleration process.
The present invention has been described in terms of the preferred embodiments thereof, and it should be understood by those skilled in the art that various modifications can be made without departing from the principles of the invention, and such modifications should also be considered as being within the scope of the invention.
Claims (6)
1. A method for controlling and distributing a compound helicopter and designing an optimal transitional route is characterized by comprising the following steps:
1) Establishing a flight mechanics model of each component, converting all forces and moments into a body shafting, and establishing a six-degree-of-freedom balance equation;
2) Determining maneuvering strategies under different flying speeds, and determining maneuvering mechanisms participating in maneuvering under different flying speeds, wherein the maneuvering strategies under different flying speeds comprise a hovering and low-speed maneuvering strategy, a high-speed forward flying maneuvering strategy and a transitional forward flying maneuvering strategy, the maneuvering mechanisms under the hovering and low-speed maneuvering strategies comprise a main rotor wing and a double-propulsion propeller, the flying speeds are 0m/s to 10m/s, the main rotor wing is subjected to collective maneuvering, longitudinal periodic variable-pitch maneuvering and transverse periodic variable-pitch maneuvering, and the double-propulsion propeller comprises an average pitch and a differential pitch; the control mechanism of the high-speed forward flight control strategy comprises a double-propulsion propeller, fixed wings and a tail wing, wherein the flight speed interval is more than 45m/s, the fixed wings are an elevator, an aileron, a rudder, a pitch channel and a pitching channel, and the double-propulsion propeller comprises an average pitch; the control mechanism of the transition front flight control strategy comprises a double-propulsion propeller, a main rotor wing, a fixed wing and a tail wing, wherein the speed interval of the transition stage is 10m/s-45m/s, the fixed wing is an elevator, an aileron, a rudder, a pitch channel and a pitching channel, the main rotor wing is subjected to collective pitch control, longitudinal periodic pitch control and transverse periodic pitch control, and the double-propulsion propeller comprises an average pitch and a differential pitch;
3) Designing a redundant control distribution coefficient taking the speed as a reference quantity, judging whether the pitch angle meets boundary conditions, and optimizing to obtain the pitch angle in each state to obtain a transition route of the compound helicopter;
4) Determining an optimization strategy of different flight speed stages according to different maneuvering strategies:
4.1 An optimization strategy of the transitional fly-forward manipulation strategy;
the pitch angle is optimized for the first time by using a Newton method, the distribution coefficient of the optimal maneuvering efficiency is used as the initial value of the distribution coefficient, one distribution coefficient corresponding to each flying speed is used for maximizing the current maneuvering efficiency, and the maneuvering coefficient is as follows:
wherein the method comprises the steps of
Expressed as pitching moment caused by the unit steering amount, < >>
Respectively a unit longitudinal period variable pitch operation and a unit elevator operation, K' cyc 、K′ ail Expressed as a longitudinal period change distance and elevator distribution coefficient in the current state, and K 'is regulated' cyc +K′ ail ＝1；
Longitudinal period change is delta lon_cyc ＝K′ cyc *δ lon The elevator steering amount is delta lon_ail ＝K′ ail *δ lon ，δ lon Is pitch channel manipulation amount;
when the distribution coefficient is optimized to zero, the control quantity is continuously reduced along with the increase of the flying speed, and the control quantity cannot meet the boundary condition through optimizing the distribution coefficient, so that the pitch angle is optimized, and the optimal power is achieved on the premise that the pitch angle meets the boundary condition;
4.2 An optimization strategy for hover and low speed maneuver strategies;
because the flying speed is low, the air flow speed acting on the wing is too low, the aerodynamic force of each control surface is small, and the efficacy of the control surface is very low, so that the distribution coefficient between the rotor wing and the wing control surface in the low-speed stage is fixed to be 1;
in the speed range of 0-10m/s, keeping the pitching attitude to be zero, avoiding the negative pitch, and the pitch angle constraint condition of the low-speed stage is as follows:
AVER(θ)-DIFF(θ)＞0(rad)
wherein AVER and DIFF are respectively the average pitch and the differential pitch in the trim state;
4.3 An optimization strategy of the high-speed fly-forward manipulation strategy;
the wing bears 90% of lift force in the high-speed stage, and the control surface has enough control efficacy to perform attitude operation, so that the distribution coefficient between the rotor wing in the high-speed stage and the wing control surface is a fixed value;
the high-speed stage ensures that the total distance is positive, and the constraint condition of the total angle of the high-speed stage is as follows:
Col＞0(rad)
where Col is the total angle in trim state, respectively;
5) Obtaining control distribution coefficients among different control mechanisms and a pitching attitude angle transition route in the transition process;
6) The calculation method is applied to the composite helicopter aircraft for calculation.
2. The method for combined helicopter maneuvering distribution and optimal transitional route design according to claim 1, wherein the method comprises the following steps: the forces and moments include rotor aerodynamic forces, propulsive propeller aerodynamic forces, wing, vertical tail, horizontal tail aerodynamic forces.
3. The method for combined helicopter maneuvering distribution and optimal transitional route design according to claim 2, wherein the method comprises the following steps: the aerodynamic force of the rotor wing is calculated as follows: adopting phyllin theory modeling, carrying out iterative solution on uniform inflow induction speed, and calculating aerodynamic force of blade micro-segments according to the phyllin theory, wherein the blade only considers first-order rigid flapping:
wherein the method comprises the steps of
Is the tangential and normal airflow velocity of the regularized blade profile, c is the blade chord length.
4. The method for combined helicopter maneuvering distribution and optimal transitional route design according to claim 2, wherein the method comprises the following steps: the aerodynamic force of the propulsion propeller is calculated as follows: dividing the blade into a plurality of micro sections, and calculating the speed at the micro sections of the blade, wherein the speed at the 1/4 chord line under the blade shafting is as follows:
v tan 、v n is the tangential velocity and the vertical velocity of the blade micro-segment, theta twist Is the negative torque of the blade; the lift-drag coefficient and the corresponding force and moment are obtained by a table look-up mode, and then the force and moment of the whole propeller to the center of the hub can be obtained by integrating.
5. The method for combined helicopter maneuvering distribution and optimal transitional route design according to claim 2, wherein the method comprises the following steps: the wing, vertical tail and horizontal tail aerodynamic forces are calculated as follows:
in a hovering state, the rotor wake induces aerodynamic loads to the wing, so that low-speed performance is reduced, and the speed of the wing micro-section under the body shafting is as follows:
the wing aerodynamic force has the following structure:
the current wing model uses the product of the hypothetical empirical coefficient and the main rotor wake uniform induction speed to perturb the wing, matching the hover load empirical coefficient k measured by Felker and Light in the experimental test of V-22 wing loads, selected to be 1.5, where
Is the aircraft speed, ρ is the air density, +.>
Is the chord length of the wing, C L 、C D 、C M Respectively the lift force, the resistance and the pitching moment coefficient.
6. The method for combined helicopter maneuvering distribution and optimal transitional route design according to claim 1, wherein the method comprises the following steps: the six-degree-of-freedom equilibrium equation described in step 1) is:
wherein I is n For the inertial matrix of the aircraft, u, v, w are the velocities under the body axis, p, q, r are the angular velocities under the body axis,
is the resultant force of all pneumatic components under the body axis, and +.>
The combined moment of each pneumatic component under the body shafting, theta, phi and phi are Euler angles of the aircraft, and the combined helicopter control quantity comprises a total pitch rod control quantity, a transverse rod control quantity, a longitudinal rod control quantity, a heading rod control quantity and an average pitch rod; when balancing calculation is carried out, the resultant force of the machine body is zero, and the combined moment of the machine body is zero, so that a motion equation set of the machine body is established as follows:
∑F x -Gsinθ＝0
∑F y +Gcosθsinφ＝0
∑F x +Gcosθcosφ＝0
∑M x ＝0
∑M y ＝0
∑M z ＝0
the number of the motion equations of the airframe is 6, the to-be-trimmed quantity is 7 total distance rod operation quantity, transverse rod operation quantity, longitudinal operation quantity, heading rod operation quantity, average pitch operation quantity, pitch angle and roll angle, an additional state is specified, the dimension of the equation set is reduced, the pitch attitude angle is specified, and finally, the to-be-trimmed quantity is total distance rod operation quantity, transverse rod operation quantity, longitudinal operation quantity, heading rod operation quantity, average pitch operation quantity and roll angle.
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