Universal modeling and balancing calculation method for single-rotor composite helicopter

Abstract

The invention provides a general modeling and trimming calculation method for a single-rotor composite helicopter. The general modeling and trimming calculation method comprises the following steps: step 1, establishing a model of the single-rotor composite helicopter; and 2, solving the model established in the step 1 by adopting a general balancing calculation method of the single-rotor composite helicopter. The general modeling and trimming calculation method provided by the invention adopts an empirical rotor wake and aerodynamic interference model, does not depend on correction factors, has enough precision, has universality and accuracy, and can be used for flight characteristic evaluation of single-rotor composite helicopters with different configurations in a conceptual design stage.

Description

Technical Field

[0001] The invention relates to the field of flight dynamics, and in particular to a general modeling and balancing calculation method for a single-rotor compound helicopter. Background Art

[0002] Helicopters have been widely used in military and civilian fields due to their vertical take-off and landing capabilities and excellent low-speed flight performance. However, the maximum flight speed of helicopters is limited by the compression effect of the forward blades and the stall of the rearward blades of the main rotor, which in turn limits their flight envelope, application range, and mission execution efficiency. With the growth of civil and military needs, people are increasingly hoping to increase the cruising speed and range of helicopters without sacrificing hovering capabilities, and compound helicopters are a promising solution (reference: RA Ormiston, Revitalizing Advanced Rotorcraft Research—and the Compound Helicopter, J. Am. Helicopter Soc. 61 (2016) 1–23. https: / / doi.org / 10.4050 / JAHS.61.011001).

[0003] Compound helicopters are usually based on traditional single-rotor helicopters with tail rotors, with wings and auxiliary thrust devices added to increase flight speed. When hovering and flying at low speeds, compound helicopters behave similarly to conventional helicopters, with the main rotor providing lift and maneuvering through cyclic pitch change. In high-speed flight, the wings and auxiliary thrust devices provide additional lift and thrust to reduce the load on the main rotor and overcome aerodynamic drag, while introducing additional control surfaces to enhance high-speed flight maneuverability. Compared with tilt-rotor aircraft, compound helicopters have hovering and low-speed performance that is closer to traditional helicopters. At the same time, they have a simple structure, can be reused from traditional helicopter production lines, and are economical. Therefore, compound helicopters have attracted widespread attention worldwide in recent years (reference: H. Yeo, Design and aeromechanics investigation of compound helicopters, Aerosp. Sci. Technol. 88 (2019) 158–173. https: / / doi.org / 10.1016 / j.ast.2019.03.010).

[0004] For novel compound helicopter configurations, it is crucial to conduct a preliminary evaluation of their flight characteristics during the conceptual design phase of development, which requires the establishment of a more reliable flight dynamics model. However, the additional wings and propellers will lead to complex aerodynamic interference problems, and compound helicopters have a variety of configurations and variable aerodynamic interference, which makes their flight dynamics modeling very difficult and less versatile. At the same time, in order to ensure the control effectiveness at low and high speeds, compound helicopters have control surfaces for both helicopters and fixed-wing aircraft, which results in redundant control problems, which means that the number of unknown quantities in balancing is greater than the number of equations, resulting in the inability of the traditional helicopter balancing calculation method to obtain a unique balancing solution, which is also not conducive to flight characteristics evaluation.

[0005] At present, many single-rotor compound helicopter configurations have been proposed, including AH-56A, X-49A, X3 and RACER. During their development, most of them use flight dynamics modeling methods based on wind tunnel test data, which is not only inefficient but also difficult to modify and only applicable to a single model. Although there are two general flight dynamics models suitable for single-rotor compound helicopters in the world, namely the HeliUM model developed by the University of Maryland and the GenHel model developed by Sikorsky, they are difficult to obtain through public channels.

[0006] The key to the flight dynamics modeling of a single-rotor compound helicopter lies in aerodynamic interference. Existing research is devoted to using computational fluid dynamics (CFD) methods to study the aerodynamic interference between the main rotor, propeller and wing, and to reveal the causes and influence of aerodynamic interference from a mechanism level (references: T. Stokkermans, M. Voskuijl, L. Veldhuis, B. Soemarwoto, Aerodynamic Installation Effects of Lateral Rotors on a Novel Compound Helicopter Configuration, in: Proc. AHS Int. 74th Annu. Forum Technol. Disp., Phoenix, Arizona, USA, 2018; F. Frey, J. Thiemeier, C. M. Keβler,E. Aerodynamic Interactions on Airbus Helicopters' Compound Helicopter RACER in Cruise Flight, J. Am. Helicopter Soc. 65 (2020) 1–14. https: / / doi.org / 10.4050 / JAHS.65.042001). However, when modeling aerodynamic interactions, empirical methods based on correction factors are still mainly used, which reduces the need for wind tunnel tests but still requires long CFD calculations. Some scholars have established a rotor-wing aerodynamic interference modeling method that is independent of the correction factor and a downwash interference velocity modeling method based on a simple rotor wake (references: PB Ferguson, A mathematical model for real time flights simulation of a generic tilt-rotor aircraft, Systems Technology INC, Mountain View, California, 1988; PB Harendra, MJ Joglekar, TMGaffey, A mathematical model for real time flight simulation of the Bell Model 301 tilt rotor research aircraft, Bell Helicopter Company, Fort Worth, Texas, 1973). These methods have been proven to be able to reflect the main characteristics of aerodynamic interference, but have not yet been combined with the flight dynamics modeling of single-rotor compound helicopters.In addition, most of the existing trim calculation methods under redundant control are targeted at a specific model (references: Y.Su, Z.Wang, Y.Cao, A hybrid trim strategy for coaxial compound helicopter, Proc.Inst.Mech.Eng.Part G J.Aerosp.Eng.237(2023)452–466.https: / / doi.org / 10.1177 / 09544100221103021; ​​Y.Yuan, R.Chen, P.Li, Trim investigation for coaxialrigid rotor helicopters using an improved aerodynamic interference model, Aerosp.Sci.Technol.85(2019)293–304.https: / / doi.org / 10.1016 / j.ast.2018.11.044), and there is a lack of a universal trim method applicable to different single-rotor compound helicopter models.

[0007] In the prior art, the flight dynamics models of single-rotor compound helicopters mostly use wind tunnel test data or correction factors obtained by CFD calculation to model aerodynamic interference, which leads to high cost, low efficiency, difficulty in modification and lack of availability. At the same time, the balancing calculation method under redundant control also lacks universality, which is not conducive to the flight characteristics evaluation in the conceptual design stage. Summary of the invention

[0008] Purpose of the invention: The technical problem to be solved by the present invention is to establish a universal flight dynamics model suitable for flight characteristics evaluation in the conceptual design stage and a balancing calculation method under redundant control in order to address the problems of high cost and poor versatility of existing single-rotor compound helicopter flight dynamics models and balancing calculation methods.

[0009] The present invention provides a general modeling and balancing calculation method for a single-rotor compound helicopter, specifically a flight dynamics modeling method consisting of independent models of various aerodynamic components and a theoretical aerodynamic interference model, and a balancing calculation method under redundant control pre-allocation, which is suitable for flight characteristic evaluation of a single-rotor compound helicopter in a conceptual design stage.

[0010] The method of the present invention comprises the following steps:

[0011] Step 1, establishing a model of a single-rotor compound helicopter;

[0012] Step 2, using the general trim calculation method of the single-rotor compound helicopter, solve the model established in step 1.

[0013] Step 1 comprises: the model of the single-rotor compound helicopter comprises a helicopter rigid body dynamics model, a main rotor aerodynamic-flapping dynamics model, a main rotor wake interference model, a main rotor wake interference model, a propeller aerodynamic model, a wing aerodynamic model under main rotor interference, a body aerodynamic model and a tail aerodynamic model;

[0014] The helicopter rigid body dynamics model is expressed as:

[0015]

[0016] Among them, V={u,v,w} T and ω={p,q,r} T are the translational velocity and angular velocity vectors of the helicopter, u, v, w are the translational velocities along the x-axis, y-axis and z-axis of the helicopter body axis system, p, q, r are the angular velocities along the x-axis, y-axis and z-axis of the helicopter body axis system, and They represent the derivative of the helicopter's translational velocity vector and angular velocity vector respectively; m and I are the helicopter's mass and inertia matrix respectively; g is the gravity vector; F * and M * Respectively represent the force vector and torque vector of *, *∈{mr,fuslg,pr,wing,tail}, mr,fuslg,pr,wing,tail represent the main rotor, fuselage, propeller, wing and tail respectively; n pr 、n wing and n tail They are the number of propellers, the number of wings and the number of tail fins.

[0017] In step 1, the main rotor aerodynamic-flapping dynamics model includes:

[0018] The blade tip trajectory plane model is used to describe the rotor flapping motion. The blade flapping angle at azimuth angle ψ and time t is denoted as β(ψ,t). The first-order Fourier series and time-varying coefficients are used to approximate the rotor blade flapping motion, which can be expressed as:

[0019] β(ψ,t)=a0(t)-a1(t)cosψ-b1(t)sinψ (2)

[0020] Where a0(t) is the rotor taper angle; a1(t) and b1(t) represent the longitudinal and lateral first-order harmonic coefficients, respectively. a1(t) and b1(t) correspond to the backward tilt angle and left tilt angle of the blade tip trajectory plane, respectively, at the physical level.

[0021] The rotor flapping dynamics equation is expressed by a0(t), a1(t), and b1(t):

[0022]

[0023] Among them, D TPP , K TPP 、f TPP They are the damping matrix, stiffness matrix and right-hand side terms of the rotor flapping dynamics equation; is the first-order derivative of a0, a1, b1, is the second-order derivative of a0, a1, b1;

[0024] The Pitt-Peters dynamic inflow theory is used to describe the main rotor inflow distribution, with the azimuth angle ψ, time t and dimensionless spanwise position recorded. The vertical inflow at Approximated by first-order Fourier series and time-varying coefficients, it can be expressed as:

[0025]

[0026] Where v0(t) is the uniform inflow term of the rotor; v 1s (t) and v 1c (t) represent the sine first-order harmonic inflow coefficient and the cosine first-order harmonic inflow coefficient respectively;

[0027] The rotor flapping dynamics equation uses v0(t), v 1s (t), v 1c (t) is expressed as:

[0028]

[0029] Where M and L are the mass matrix and inflow gain matrix respectively; C T , C L , C M They represent the aerodynamic lift, rolling moment and pitching moment of the main rotor respectively; is v0, v 1s 、v 1c The first derivative of ;

[0030] The force vector F of the main rotor is obtained by integrating the aerodynamic force and inertial force during the rotor rotation cycle: mr and the moment vector M mr .

[0031] In step 1, the main rotor wake interference model includes: the rotor wake model is defined in the blade tip trajectory plane coordinate system, and the X axis of the blade tip trajectory plane coordinate system is w The direction is consistent with that of the free flow, and the angle between it and the free flow is -α mr , the Z axis is perpendicular to the blade tip trajectory plane and points downward, while the Y axis is determined according to the right-hand rule; the evolution of the wake is calculated using the pseudo-time τ wIt means that only the component v of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane is considered iw (τ w ), then the z and x positions of the rotor wake centerline in the blade tip trajectory plane coordinate system are defined by derivatives:

[0032]

[0033] Where d represents the differential; x w 、z w They represent the x-axis position and z-axis position of the rotor wake centerline, Respectively represent x w 、z w The first derivative of ;

[0034] When τ w = 0, the wake induced airflow velocity is equal to the uniform inflow size v0 at the main rotor blade disc;

[0035] The rotor wake induced velocity is assumed to increase linearly with the evolution time, and the component v of the wake induced airflow velocity perpendicular to the blade tip trajectory plane is iw (τ w ) and the z-axis position z of the rotor wake centerline w (τ w ) and the x-axis position x w (τ w ) is expressed as:

[0036] v iw (τ w )=A1τ w +v0 (7)

[0037]

[0038] Among them, A1 is v iw (τ w ) with respect to the time derivative of the wake evolution is obtained through the following rotor wake evolution equation:

[0039]

[0040] Among them, τ w,far is the evolution time corresponding to the rotor wake evolving to a distance 1.5 times the radius; R mr is the main rotor radius;

[0041] The rotor wake induced airflow velocity is set to be uniformly distributed on the circular cross section, and the rotor wake cross section is set to remain circular during the evolution process. The rotor wake radius R w According to the continuity equation of momentum theory, we get:

[0042]

[0043] Based on formula (7), formula (8) and formula (10), we can get r at any position pos The rotor wake induced airflow velocity v at pos , v pos Represented as r pos The nonlinear function is denoted by v pos =v pos (r pos ), the specific expression is:

[0044] v pos =v pos (r pos )

[0045] Given

[0046] but

[0047] like

[0048] v pos =(A1τ w,pos +v0)e3 TPP

[0049] otherwise

[0050] v pos =0

[0051] Among them, e k TPP Represents the unit vector of the blade tip trajectory plane coordinate system, k = 1, 2, 3; Yes pos Position coordinates in the blade tip trajectory plane coordinate system, j = x, y, z.

[0052] In step 1, the propeller aerodynamic model includes:

[0053] The propeller flow velocity is expressed as:

[0054] v pr =V+ω×r cg2pr -v mr2pr =μ pr Ω pr R pr e1 PRW +0e2 PRW +λ pr Ω pr R pr e3 PRW (12)

[0055] Among them, vpr is the propeller flow velocity vector; r cg2pr represents the position vector from the center of gravity to the propeller; v mr2pr is the induced velocity of the rotor wake at the center of the propeller; e k PRW are the three unit vectors of the propeller wind axis system, k = 1, 2, 3, the XOY plane is parallel to the propeller hub, and the X axis is facing the incoming flow direction; μ pr and λ pr are the propeller advance ratio and inflow ratio respectively; Ω pr and R pr are the speed and radius of the propeller respectively;

[0056] The propeller inflow velocity is obtained through momentum theory, and the equation is:

[0057]

[0058] Among them, C T,pr is the propeller thrust coefficient; v i,pr is the propeller induced speed;

[0059] Propeller force vector F mr and the moment vector M mr Only tensile forces and reactive torques are considered.

[0060] In step 1, the wing aerodynamic model under the interference of the main rotor includes: the wing is divided into a wake area and a non-wake area;

[0061] In order to calculate the wake area and non-wake area, the wing is divided into small areas using a grid. The total number of areas is N. wing , the i-th region and its center point are respectively denoted as s grid,i and r mr2wing,i , the rotor aerodynamic disturbance v in the i-th region is obtained by formula (11): mr2wing,i ; Define the set of areas affected by the rotor The total area of ​​the wing disturbed by the rotor wake is S wing,inMR and the average air velocity v in the area wing,inMR It can be expressed approximately as:

[0062]

[0063] Among them, r cg 2wi ng, i represents the position vector from the center of gravity to the center point of the i-th region;

[0064] Then the wing aerodynamic force F in the main rotor wake interference area is wing,inMR It is expressed as:

[0065] F wing,inMR =q wing,inMR S wing,inMR C wing (16)

[0066] Among them, q wing,inMR It is dynamic pressure; C wing is the aerodynamic coefficient of the wing;

[0067] Get the wing aerodynamic force F outside the main rotor wake interference area wing,outMR , the total aerodynamic load of the wing is expressed as F wing,inMR With F wing,outMR sum.

[0068] In step 1, the aerodynamic model of the fuselage includes: the incoming flow velocity v of the fuselage fuslg It is expressed as:

[0069]

[0070] Among them, v mr2fuslg is the main rotor wake induced speed at the aerodynamic center of the fuselage; e k F Represents the unit vector of the fuselage coordinate system. The X, Y, and Z axes of the fuselage coordinate system point to the front, right, and bottom of the fuselage respectively. k = 1, 2, 3; are the velocity components corresponding to the X, Y, and Z axes, j = x, y, z;

[0071] Aerodynamic force F of the fuselage fuslg and moment M fuslg It is expressed as:

[0072]

[0073]

[0074] Among them, e k FW is the unit vector of the fuselage wind coordinate system, the X-axis of the fuselage wind coordinate system is facing the incoming flow direction, k = 1, 2, 3; q fuslg It is dynamic pressure; S fuslg and L fuslg are the equivalent area and length of the fuselage respectively; are the aerodynamic coefficients of *, *=Drag, Side, Lift, Roll, Pitch, Yaw, Drag, Side, Lift, Roll, Pitch, Yaw represent drag, lateral force, lift, rolling moment, pitching moment and yaw moment respectively.

[0075] In step 1, the tail aerodynamic model includes: considering the rotor aerodynamic interference, according to the tail aerodynamic center position r mr2tail Use formula (11) to calculate the disturbance airflow velocity v mr2tail , and then superimpose the free flow to get the tail relative flow velocity, then the tail aerodynamic force F tail It is expressed as:

[0076] F tail =q tail S tail C tail (19)

[0077] Among them, q tail is the dynamic pressure; S tail is the tail area; C tail is the aerodynamic coefficient.

[0078] Step 2 includes: The control surfaces of the single-rotor compound helicopter include: the total pitch of the main rotor δ mr,col and longitudinal pitch control δ mr,lon , lateral pitch controlδ mr,lat 、N pr Collective pitch control of each propeller N wing Control surfaces on each wing and N tail Tail rudder control Total N tot =N pr +N wing +N tail +3 control surfaces, plus helicopter pitch and roll attitude, totaling N trim =N pr +N wing +N tail +5 unknown balance quantities;

[0079] The allocation method of each control surface is established according to the channels and functions affected by different control surfaces. The obtained control allocation strategy is expressed as:

[0080]

[0081] Among them, δ in is the input column vector of the five manipulation channels, δ c is the control surface column vector of the single-rotor compound helicopter, [λ] is the control allocation coefficient matrix from the control surface column vector to the control channel input column vector; δ ver , δ roll , δ pitch , δ yaw , δ thrustThey represent the vertical control channel, the roll control channel, the pitch control channel, the yaw control channel and the thrust control channel respectively; They represent the control distribution coefficients of the vertical, lateral and longitudinal control channels of the main rotor respectively; is the lateral control channel allocation coefficient of the wing; They represent the distribution coefficients of the propeller's yaw and thrust control channels respectively; They represent the distribution coefficients of the vertical, longitudinal and yaw control channels of the tail wing respectively;

[0082] Based on the control allocation strategy, a trim calculation method is adopted to calculate the trim state of the single-rotor compound helicopter.

[0083] In step 2, the trim state of the single-rotor compound helicopter is calculated using a trim calculation method, specifically including:

[0084] Step 2-1, set the initial value of the control channel input and the helicopter roll angle, input the flight speed corresponding to the trim state into the control allocation strategy and the helicopter pitch angle change curve, and obtain the control allocation coefficient matrix and the preset helicopter pitch angle at the trim speed point;

[0085] Step 2-2, invert the control distribution coefficient matrix to obtain the control surface deflection corresponding to the current control channel input, input the control surface deflection and the preset helicopter pitch angle and helicopter roll angle initial values ​​into formula (1), and obtain the acceleration and angular acceleration corresponding to the initial values;

[0086] Step 2-3, if the acceleration and angular acceleration are zero, the balancing is completed, otherwise the balancing amount is adjusted by the gradient descent method until the balancing is completed.

[0087] Beneficial effect: The present invention adopts the above technical scheme compared with the prior art. Most of the prior art is based on correction factors obtained by wind tunnel test data or CFD calculation, which are costly, inefficient, difficult to modify and difficult to obtain, and the balancing calculation method also lacks universality. The general modeling and balancing calculation method proposed in the present invention adopts an empirical rotor wake and aerodynamic interference model, is independent of the correction factor and has sufficient accuracy, and has universality and accuracy, and can be used for the flight characteristics evaluation of single-rotor compound helicopters of different configurations in the conceptual design stage. BRIEF DESCRIPTION OF THE DRAWINGS

[0088] The present invention will be further described in detail below in conjunction with the accompanying drawings and specific embodiments, and the above and / or other advantages of the present invention will become more clear.

[0089] Figure 1 It is a schematic diagram of the general modeling and balancing calculation method of a single-rotor compound helicopter.

[0090] Figure 2 It is a schematic diagram of the simplified rotor wake model.

[0091] Figure 3 It is a schematic diagram of the empirical rotor-wing aerodynamic interference model.

[0092] Figure 4 It is a schematic diagram of the trim calculation method for a single-rotor compound helicopter.

[0093] Figure 5 It is a schematic diagram of the wind tunnel trim test of the 12kg twin-propeller propulsion compound helicopter prototype.

[0094] Figure 6a , Figure 6b , Figure 6c , Figure 6d , Figure 6e It is a schematic diagram comparing the model calculation results with the wind tunnel balancing test results. DETAILED DESCRIPTION

[0095] This embodiment provides a general modeling and trim calculation method for a single-rotor compound helicopter, which is applicable to the flight characteristics evaluation of single-rotor compound helicopters of different configurations in the conceptual design stage. Figure 1 The figure shows a schematic diagram of the general modeling and trim calculation method of a single-rotor compound helicopter. The upper part shows the motion equation of the single-rotor compound helicopter, and the lower part shows the method for reducing the number of unknown quantities. Ultimately, the number of unknown quantities is matched with the number of motion equations, so that a unique solution to the trim state can be obtained.

[0096] Firstly, a general flight dynamics model of a single-rotor compound helicopter is established, which consists of the main rotor aerodynamic-flapping dynamics model, the helicopter rigid body dynamics model, the propeller aerodynamic model, the fuselage aerodynamic model, the wing aerodynamic model, the tail aerodynamic model and the aerodynamic interference model. The overall motion equation can be expressed using the rigid body dynamics model.

[0097] The main rotor aerodynamic-flapping dynamics model uses the blade tip trajectory plane model to describe the rotor flapping motion, and uses the Pitt-Peters dynamic inflow theory to describe the main rotor inflow distribution. The force vector F of the main rotor is obtained by integrating the aerodynamic force and inertial force during the rotor rotation cycle. mr and the moment vector M mr Since all aerodynamic components of a single-rotor compound helicopter are located below the main rotor, the downwash of the main rotor has a significant impact on these components during hovering and low-speed flight. Figure 2 The simplified rotor wake model shown in the figure is combined with the Pitt-Peters dynamic inflow model and the blade tip trajectory plane dynamics model to obtain the pos The rotor wake induced airflow velocity v at pos, used to consider the aerodynamic interference between the main rotor and other components.

[0098] The propeller aerodynamic model is similar to that of the main rotor, but the dynamic changes of flapping and inflow are ignored. The propeller inflow velocity is first expressed as the superposition of the propeller relative inflow velocity and the induced velocity of the rotor wake at the center of the propeller, and then the propeller inflow velocity is obtained through momentum theory. Finally, the propeller thrust and anti-torque are obtained through the calculation method of the propeller consistent with the main rotor.

[0099] In order to consider the significant influence of the main rotor wake on the wing, the present invention proposes an empirical rotor-wing aerodynamic interference model without correction factors, so that it can be applicable to the modeling of various single-rotor compound helicopters. Figure 3 The empirical rotor-wing aerodynamic interference model is shown. The wing is divided into a gray "wake area" and a white "non-wake area". The downwash of the main rotor mainly affects the "wake area" and has no effect on the "non-wake area". The areas of the "wake area" and "non-wake area" are calculated by dividing the wing into small areas using a grid. Therefore, the wing aerodynamic force in the main rotor wake interference area can be expressed as the superposition of the wing aerodynamic forces inside and outside the main rotor wake interference area.

[0100] The aerodynamic model of the fuselage uses the aerodynamic coefficient interpolation method. Whether to consider the aerodynamic interference of the rotor on the fuselage is determined according to whether the aerodynamic center of the fuselage is located in the rotor wake. The aerodynamic interference speed is determined by the rotor wake induced airflow speed at the aerodynamic center of the fuselage.

[0101] The tail aerodynamic model uses the aerodynamic coefficient interpolation method. Whether to consider the aerodynamic interference of the rotor on the tail is determined according to whether the aerodynamic center of the tail is located in the rotor wake. The aerodynamic interference speed is determined by the rotor wake induced airflow speed at the aerodynamic center of the tail.

[0102] Then, a general trim calculation method suitable for single-rotor compound helicopters is established (e.g. Figure 4 The control surfaces of single-rotor compound helicopters are classified into: main rotor collective pitch (δ mr,col ) and longitudinal and lateral pitch control (δ mr,lon and δ mr,lat ), N pr Collective pitch control of each propeller N wing Control surfaces on each wing and N tail Tail rudder control Total N tot =N pr +N wing +N tail +3 control surfaces, plus helicopter pitch and roll attitude, totaling N trim =N pr+N wing +N tail +5 trim unknowns. By introducing the control distribution coefficient and pre-designing the pitch angle variation curve with forward flight speed, the number of trim unknowns is reduced to 6, which is equal to the number of 6 equations that the rigid body dynamics equation can provide, thus obtaining a unique trim solution.

[0103] Finally, a wind tunnel trim test was carried out on a twin-propeller propulsion compound helicopter based on a 12 kg prototype (e.g. Figure 5 shown). Figure 6a , Figure 6b , Figure 6c , Figure 6d , Figure 6e The model calculation results in the experiment are in good agreement with the experimental values, which proves the rationality of the method in this embodiment.

[0104] In a specific embodiment of the present invention, a twin-propeller propulsion compound helicopter is used as the object, and a wind tunnel balancing test is carried out based on a 12kg prototype, which shows that the method of this embodiment can reasonably predict the flight characteristics of a single-rotor compound helicopter.

[0105] like Figure 1 The figure shows a schematic diagram of the general modeling and trim calculation method of a single-rotor compound helicopter. The upper part shows the motion equation of the single-rotor compound helicopter, and the lower part shows the method for reducing the number of unknown quantities. Ultimately, the number of unknown quantities is matched with the number of motion equations, so that a unique solution to the trim state can be obtained.

[0106] 1. General modeling method for single-rotor compound helicopter

[0107] The single-rotor compound helicopter model consists of a helicopter rigid body dynamics model, a main rotor aerodynamic-flapping dynamics model, a main rotor wake interference model, a main rotor wake interference model, a propeller aerodynamic model, a wing aerodynamic model under main rotor interference, a fuselage aerodynamic model, and a tail aerodynamic model. Since all aerodynamic components of a single-rotor compound helicopter are located under the main rotor, the downwash of the main rotor has a significant impact on these components during hovering and low-speed flight. However, accurate consideration of these aerodynamic interactions usually requires expensive and time-consuming wind tunnel experiments to obtain correction factors. In the conceptual design development stage, generalized and simplified considerations of these interactions are more conducive to the preliminary evaluation of design options. The following introduces a single-rotor compound helicopter flight dynamics modeling method that does not rely on correction factors.

[0108] The overall motion equation of a single-rotor compound helicopter can be expressed using a rigid body dynamics model as follows (ignoring the aerodynamic torque of the wing and tail):

[0109]

[0110] Among them, V={u,v,w} T and ω={p,q,r} T are the helicopter's translational velocity and angular velocity vectors, respectively; m and I are the helicopter's mass and inertia matrix, respectively; g is the gravity vector; F * and M * Respectively represent the force vector and torque vector of *, *∈{mr,fuslg,pr,wing,tail}, mr,fuslg,pr,wing,tail represent the main rotor, fuselage, propeller, wing and tail respectively; n pr 、n wing and n tail These are the number of propellers, wings and tail fins respectively.

[0111] The specific modeling methods of the main rotor aerodynamic-flapping dynamics model, the main rotor wake interference model, the main rotor wake interference model, the propeller aerodynamic model, the wing aerodynamic model under the main rotor interference, the fuselage aerodynamic model and the tail aerodynamic model are as follows:

[0112] 1.1 Main rotor aerodynamic-flapping dynamics model

[0113] The main rotor provides the main lift during hovering and low-speed flight, and the blade tip trajectory plane (TPP) model is used to describe the rotor flapping motion. The blade flapping angle at azimuth angle ψ and time t is β(ψ,t), which can be approximated by the first-order Fourier series and time-varying coefficients as follows:

[0114] β(ψ,t)=a0(t)-a1(t)cosψ-b1(t)sinψ (2)

[0115] Where a0(t) is the rotor taper angle; a1(t) and b1(t) represent the longitudinal and lateral first-order harmonic coefficients, respectively, which correspond to the backward and left inclination angles of the blade tip trajectory plane at the physical level.

[0116] The rotor flapping dynamics equation can be expressed by a0(t), a1(t), and b1(t) as follows:

[0117]

[0118] Among them, D TPP , K TPP 、f TPPThey are the damping matrix, stiffness matrix and right-hand side term of the rotor flapping dynamics equation. Detailed expressions can be found in the reference: RTNChen, A Simplified Rotor System Mathematical Model for piloted flight dynamics simulation, Ames Research Center, Moffett Field, California, 1979.

[0119] The Pitt-Peters dynamic inflow theory is used to describe the main rotor inflow distribution, with the azimuth angle ψ, time t and dimensionless spanwise position recorded. The vertical inflow at It can be approximated using a first-order Fourier series and time-varying coefficients as follows:

[0120]

[0121] Where v0(t) is the uniform inflow term of the rotor; v 1s (t) and v 1c (t) represent the sine first-order harmonic inflow coefficient and cosine first-order harmonic inflow coefficient respectively.

[0122] The rotor flapping dynamics equation can be expressed as v0(t), v 1s (t), v 1c (t) is expressed as follows:

[0123]

[0124] Where M and L are the mass matrix and inflow gain matrix respectively; C T , C L , C M is the aerodynamic lift, rolling moment and pitching moment of the main rotor. Detailed expressions can be found in the reference: DMPitt, D.A.Peters, Theoretical Prediction of Dynamic-Inflow Derivatives, in: Proc. 6th Eur. Rotorcr. Powered Lift Aircr. Forum, Bristol, England, 1980.

[0125] The force vector F of the main rotor is obtained by integrating the aerodynamic force and inertial force during the rotor rotation cycle: mr and the moment vector M mr, for detailed expressions, see the reference: HB Hilbert, A Mathematical Model of the UH-60 Helicopter, Ames Research Center, Moffett Field, California, 1984.

[0126] 1.2 Main rotor wake interference model

[0127] The present invention proposes Figure 2 The simplified rotor wake model shown is combined with the Pitt-Peters dynamic inflow model and the tip track plane dynamics model to consider the aerodynamic interference between the main rotor and other components.

[0128] The rotor wake model is defined in the blade tip trajectory plane coordinate system, with its X axis and free flow V w The direction is consistent with that of the free flow, and the angle between it and the free flow is -α mr , the Z axis is perpendicular to the blade tip trajectory plane and the Y axis is determined by the right-hand rule. The evolution of the wake is calculated using the pseudo-time τ w It means that only the component of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane is considered, denoted as v iw (τ w ), then the z and x positions of the rotor wake centerline in the blade tip trajectory plane coordinate system can be defined by their derivatives as follows:

[0129]

[0130] When τ w = 0, the wake induced airflow velocity is equal to the uniform inflow size v0 at the main rotor disc. According to momentum theory, the wake induced airflow velocity increases from v0 to 2v0 at infinity. In practical applications, it is reasonable to assume that the wake is fully developed at a distance of 1.5 times the radius of the main rotor (reference: G. Guglieri, P. Marguerettaz, Dynamic Stability of a Helicopter with an External Suspended Load, J. Am. Helicopter Soc. 59 (2014) 1–12. https: / / doi.org / 10.4050 / JAHS.59.042002). Assuming that the rotor wake induced velocity increases linearly with the evolution time, v iw (τ w ), z w (τ w ), x w (τ w ) is expressed as follows:

[0131] v iw (τ w )=A1τ w +v0 (7)

[0132]

[0133] Where A1 is v iw (τ w ) with respect to the time derivative of the wake evolution is obtained through the following rotor wake evolution equation:

[0134]

[0135] In the formula, τ w,far It is the evolution time corresponding to the rotor wake evolving to a distance 1.5 times the radius.

[0136] Assuming that the rotor wake induced airflow velocity is uniformly distributed on the circular cross section, and assuming that its cross section remains circular during the evolution process, its radius can be obtained according to the continuity equation of momentum theory:

[0137]

[0138] Based on formula (7), formula (8) and formula (10), we can get r at any position pos The rotor wake induced airflow velocity v at pos . pos Represented as r pos The nonlinear function is denoted by v pos =v pos (r pos ), the specific expression is:

[0139] v pos =v pos (r pos )

[0140] Given

[0141] but

[0142] like

[0143] v pos =(A1τ w,pos +v0)e3 TPP

[0144] otherwise

[0145] v pos =0

[0146] Among them, e kTPP Represents the unit vector of the blade tip trajectory plane coordinate system, k = 1, 2, 3; Yes pos Position coordinates in the blade tip trajectory plane coordinate system, j = x, y, z.

[0147] 1.3 Propeller aerodynamic model

[0148] The propeller can provide additional thrust for the single-rotor compound helicopter, offset the torque of the main rotor and provide additional control surfaces. Its aerodynamic model is similar to that of the main rotor, but the dynamic changes of flapping and inflow are ignored. First, the propeller flow velocity is expressed as:

[0149] v pr =V+ω×r cg2pr -v mr2pr =μ pr Ω pr R pr e1 PRW +0e2 PRW +λ pr Ω pr R pr e3 PRW (12)

[0150] Among them, v pr is the propeller flow velocity vector; r cg2pr represents the position vector from the center of gravity to the propeller; v mr2pr is the induced velocity of the rotor wake at the center of the propeller; e k PRW are the three unit vectors of the propeller wind axis system, k = 1, 2, 3, the XOY plane is parallel to the propeller hub, and the X axis is facing the incoming flow direction; μ pr and λ pr are the propeller advance ratio and inflow ratio respectively; Ω pr and R pr are the speed and radius of the propeller respectively;

[0151] The propeller inflow velocity is obtained through momentum theory, and the equation is as follows:

[0152]

[0153] In the formula, C T,pr is the propeller thrust coefficient.

[0154] Propeller force vector F mr and the moment vector M mr Only the thrust and counter-torque are considered, and the calculation method is the same as that of the main rotor.

[0155] 1.4 Aerodynamic model of the wing under the interference of the main rotor

[0156] In hovering and low-speed flight, the wake of the main rotor has a significant impact on the wing, so the aerodynamic interference between the rotor and the wing needs to be considered. Referring to the rotor-airfoil aerodynamic interference model of the tilt-rotor aircraft (references: PB Ferguson, A mathematical model for real time flight simulation of a generic tilt-rotor aircraft, Systems Technology INC, Mountain View, California, 1988; PB Harendra, MJ Joglekar, TMGaffey, A mathematical model for real time flight simulation of the Bell Model 301 tilt rotor research aircraft, Bell Helicopter Company, Fort Worth, Texas, 1973), the present invention proposes an empirical rotor-wing aerodynamic interference model without a correction factor, so that it can be applied to the modeling of various single-rotor compound helicopters. Figure 3 The empirical rotor-wing aerodynamic interference model is shown. The wing is divided into a gray "wake area" and a white "non-wake area". The downwash of the main rotor mainly affects the "wake area" and has no effect on the "non-wake area".

[0157] In order to calculate the "wake area" and "non-wake area", the wing is divided into small areas using a grid. The total number of areas is N wing , the i-th small area and its center point are respectively recorded as s grid,i and r mr2wing,i , then the rotor aerodynamic disturbance v in the i-th region can be obtained by formula (11): mr2wing,i The set of regions affected by the rotor is defined as Then the total area of ​​the wing affected by the rotor wake and the average airflow velocity in the area can be approximately expressed as:

[0158]

[0159] In the formula, r cg2wing,i Represents the position vector from the centroid to the center point of the i-th region.

[0160] Then the wing aerodynamic force in the main rotor wake interference area can be expressed as (ignoring the aerodynamic moment of the wing):

[0161] F wing,inMR =q wing,inMR Swing,inMR C wing (16)

[0162] In the formula, q wing,inMR is the dynamic pressure; C wing is the aerodynamic coefficient of the wing.

[0163] The aerodynamic force F of the wing outside the main rotor wake interference area wing,outMR A similar method can be used to obtain the total aerodynamic load of the wing, so that the total aerodynamic load of the wing can be expressed as F wing,inMR With F wing,outMR sum.

[0164] 1.5 Aircraft aerodynamic model

[0165] The aerodynamic model of the fuselage uses the aerodynamic coefficient interpolation method. If the aerodynamic center of the fuselage is located in the rotor wake, the downwash velocity of the rotor will be superimposed on the free flow. fuslg It is expressed as:

[0166]

[0167] Among them, v mr2fuslg is the main rotor wake induced speed at the aerodynamic center of the fuselage; e k F Represents the unit vector of the fuselage coordinate system. The X, Y, and Z axes of the fuselage coordinate system point to the front, right, and bottom of the fuselage respectively. k = 1, 2, 3; are the velocity components corresponding to the X, Y, and Z axes, j = x, y, z;

[0168] The aerodynamic forces and moments on the fuselage are expressed as:

[0169]

[0170] Among them, e k FW is the unit vector of the fuselage wind coordinate system, the X-axis of the fuselage wind coordinate system is facing the incoming flow direction, k = 1, 2, 3; q fuslg is the dynamic pressure; S fuslg and L fuslg are the equivalent area and length of the fuselage respectively; are the aerodynamic coefficients of *, *=Drag, Side, Lift, Roll, Pitch, Yaw, Drag, Side, Lift, Roll, Pitch, Yaw represent drag, lateral force, lift, rolling moment, pitching moment and yaw moment respectively.

[0171] 1.6 Tail wing aerodynamic model

[0172] The tail model is similar to the wing model, but due to its smaller area, no meshing is performed when considering the rotor aerodynamic interference. Instead, the tail aerodynamic center position r mr2tail Use formula (11) to calculate the disturbance airflow velocity v mr2tail , and then superimpose the free flow to obtain the tail relative flow velocity, then the tail aerodynamic force is expressed as

[0173] F tail =q tail S tail C tail (19)

[0174] In the formula, q tail is the dynamic pressure; S tail is the tail area; C tail is the aerodynamic coefficient.

[0175] 2. General trim calculation method for single-rotor compound helicopters

[0176] The control surfaces of a single-rotor compound helicopter include: main rotor collective pitch (δ mr,co l) and longitudinal and lateral pitch control (δ mr, l on and δ mr,lat ), N pr Collective pitch control of each propeller N wing Control surfaces on each wing and N tail Tail rudder control Total N tot =N pr +N wing +N tail +3 control surfaces, plus helicopter pitch and roll attitude, totaling N trim =N pr +N wing +N tail +5 unknown balance quantities.

[0177] Since the rigid body dynamics equation can only provide 6 equations, in order to obtain a unique balancing solution, it is necessary to introduce the control distribution coefficient to reduce the number of balancing unknowns to 6. To this end, the distribution method of each control surface is established according to the channels and functions affected by different control surfaces, and the pitch angle variation curve with forward flight speed is pre-designed to ensure good flight characteristics. Table 1 lists the possible control distribution coefficients, which are all functions of the forward flight speed.

[0178] Table 1

[0179]

[0180] Since the wing control surfaces and tail control surfaces have control effectiveness only at a certain forward flight speed, the corresponding control distribution coefficients are set to zero in the hovering state, and only the control surfaces of the main rotor and propeller are considered. As the forward flight speed gradually increases, the control effectiveness of the wing control surfaces and tail control surfaces gradually increases. At this time, the control distribution coefficients of the main rotor, propeller, wing and tail are adjusted to keep the total control effectiveness uniform. In the high-speed flight state, the control of the wing and tail control surfaces and the propeller are mainly relied on to provide propulsion. At this time, the distribution coefficient of the main rotor control surface is zero. The resulting control distribution strategy is expressed as follows:

[0181]

[0182] In the formula, δ in is the input column vector of the five manipulation channels, δ c is the control surface column vector of the single-rotor compound helicopter, [λ] is the control allocation coefficient matrix from the control surface column vector to the control channel input column vector, which is a function of the forward flight speed.

[0183] Based on the above manipulation allocation strategy, Figure 4 The trim calculation method shown calculates the trim state of a single-rotor compound helicopter. The unknown trim quantities include 5 control channel input quantities and 1 helicopter roll angle. The helicopter rigid body motion equations include 3 acceleration equations and 3 angular acceleration equations. The number of unknown trim quantities and the number of equations are both 6, so a unique trim solution can be obtained.

[0184] The specific process is:

[0185] (1) Setting the initial values ​​of the control channel input and the helicopter roll angle, inputting the flight speed corresponding to the trim state into the control allocation strategy and the helicopter pitch angle change curve, and obtaining the control allocation coefficient matrix and the preset helicopter pitch angle at the trim speed point;

[0186] (2) Invert the control distribution coefficient matrix to obtain the control surface deflection corresponding to the current control channel input, input it and the preset helicopter pitch angle and helicopter roll angle initial values ​​into the helicopter rigid body motion equation to obtain the acceleration and angular acceleration corresponding to the initial values;

[0187] (3) If the acceleration and angular acceleration are zero, the balancing is completed; otherwise, the balancing amount is adjusted by the gradient descent method until the balancing is completed.

[0188] 3. Trim results for a sample single-rotor compound helicopter

[0189] In order to verify the correctness of the general modeling and balancing method of the above-mentioned single-rotor compound helicopter, the wind tunnel balancing test data of a 12kg twin-propeller propulsion compound helicopter prototype was used to verify the balancing results (such as Figure 5 Table 2 shows the prototype parameters.

[0190] Table 2

[0191]

[0192] The wind tunnel trim test was carried out in the low-speed wind tunnel of the National Key Laboratory of Helicopter Dynamics of Nanjing University of Aeronautics and Astronautics. The relative incoming flow speeds were 0m / s, 5m / s, 10m / s, 15m / s and 20m / s, and the pitch angle was 2°. During the test, the control distribution coefficients of the wing and tail control surfaces were set to zero, and the other control surfaces were adjusted so that the six force elements measured by the six-component balance were zero, and the control amount of each control surface was recorded at this time. Figure 6a , Figure 6b , Figure 6c , Figure 6d , Figure 6e It is a comparison between the model calculation results and the test values, and it can be found that the two are in good agreement, indicating that the general modeling and balancing method for the single-rotor compound helicopter established by the method of this embodiment has good accuracy and can be used for preliminary flight characteristics evaluation.

[0193] The present invention provides a general modeling and trim calculation method for a single-rotor compound helicopter. There are many methods and ways to implement the technical solution. The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principle of the present invention. These improvements and modifications should also be regarded as the protection scope of the present invention. All components not specified in this embodiment can be implemented by existing technologies.

Claims

1. A general modeling and trim calculation method for a single-rotor compound helicopter, characterized in that: The following steps are involved: Step 1, building a model of a single-rotor compound helicopter; Step 2, using the general balancing calculation method of the single-rotor compound helicopter, solve the model established in step 1.

2. The method according to claim 1, characterized in that Step 1 comprises: the model of the single-rotor compound helicopter comprises a helicopter rigid body dynamics model, a main rotor aerodynamic-flapping dynamics model, a main rotor wake interference model, a main rotor wake interference model, a propeller aerodynamic model, a wing aerodynamic model under main rotor interference, a body aerodynamic model and a tail aerodynamic model; The helicopter rigid body dynamics model is expressed as: Among them, V={u,v,w} T and ω={p,q,r} T are the translational velocity and angular velocity vectors of the helicopter, u, v, w are the translational velocities along the x-axis, y-axis and z-axis of the helicopter body axis system, p, q, r are the angular velocities along the x-axis, y-axis and z-axis of the helicopter body axis system, and They represent the derivative of the helicopter's translational velocity vector and angular velocity vector respectively; m and I are the helicopter's mass and inertia matrix respectively; g is the gravity vector; F * and M * Respectively represent the force vector and torque vector of *, *∈{mr,fuslg,pr,wing,tail}, mr,fuslg,pr,wing,tail represent the main rotor, fuselage, propeller, wing and tail respectively; n pr 、n wing and n tail They are the number of propellers, the number of wings and the number of tail fins.

3. The method according to claim 2, characterized in that In step 1, the main rotor aerodynamic-flapping dynamics model includes: The blade tip trajectory plane model is used to describe the rotor flapping motion. The blade flapping angle at azimuth angle ψ and time t is denoted as β(ψ,t). The first-order Fourier series and time-varying coefficients are used to approximate the rotor blade flapping motion, which can be expressed as: β(ψ,t)=a0(t)-a1(t)cosψ-b1(t)sinψ (2) Where a0(t) is the rotor taper angle; a1(t) and b1(t) represent the longitudinal and lateral first-order harmonic coefficients, respectively. a1(t) and b1(t) correspond to the backward tilt angle and left tilt angle of the blade tip trajectory plane, respectively, at the physical level. The rotor flapping dynamics equation is expressed by a0(t), a1(t), and b1(t): Among them, D TPP , K TPP 、f TPP They are the damping matrix, stiffness matrix and right-hand side terms of the rotor flapping dynamics equation respectively; is the first-order derivative of a0, a1, b1, is the second-order derivative of a0, a1, b1; The Pitt-Peters dynamic inflow theory is used to describe the main rotor inflow distribution, with the azimuth angle ψ, time t and dimensionless spanwise position recorded. The vertical inflow at Approximated by first-order Fourier series and time-varying coefficients, it can be expressed as: Where v0(t) is the uniform inflow term of the rotor; v 1s (t) and v 1c (t) represent the sine first-order harmonic inflow coefficient and the cosine first-order harmonic inflow coefficient respectively; The rotor flapping dynamics equation uses v0(t), v 1s (t), v 1c (t) is expressed as: Where M and L are the mass matrix and inflow gain matrix respectively; C T , C L , C M They represent the aerodynamic lift, rolling moment and pitching moment of the main rotor respectively; is v0, v 1s 、v 1c The first derivative of ; The force vector F of the main rotor is obtained by integrating the aerodynamic force and inertial force during the rotor rotation cycle: mr and the moment vector M mr .

4. The method according to claim 3, characterized in that In step 1, the main rotor wake interference model includes: the rotor wake model is defined in the blade tip trajectory plane coordinate system, and the X axis of the blade tip trajectory plane coordinate system is w The direction is consistent with that of the free flow, and the angle between it and the free flow is -α mr , the Z axis is perpendicular to the blade tip trajectory plane and points downward, while the Y axis is determined according to the right-hand rule; the evolution of the wake is calculated using the pseudo-time τ w It means that only the component v of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane is considered iw (τ w ), then the z and x positions of the rotor wake centerline in the blade tip trajectory plane coordinate system are defined by derivatives: Where d represents the differential; x w 、z w They represent the x-axis position and z-axis position of the rotor wake centerline, Respectively represent x w 、z w The first derivative of ; When τ w = 0, the wake induced airflow velocity is equal to the uniform inflow size v0 at the main rotor blade disc; The rotor wake induced velocity is assumed to increase linearly with the evolution time, and the component v of the wake induced airflow velocity perpendicular to the blade tip trajectory plane is iw (τ w ) and the z-axis position z of the rotor wake centerline w (τ w ) and the x-axis position x w (τ w ) is expressed as: v iw (t w )=A1τ w +v0 (7) Among them, A1 is v iw (τ w ) with respect to the time derivative of the wake evolution is obtained through the following rotor wake evolution equation: Among them, τ w,far is the evolution time corresponding to the rotor wake evolving to a distance 1.5 times the radius; R mr is the main rotor radius; The rotor wake induced airflow velocity is set to be uniformly distributed on the circular cross section, and the rotor wake cross section is set to remain circular during the evolution process. The rotor wake radius R w According to the continuity equation of momentum theory, we get: Based on formula (7), formula (8) and formula (10), we can get r at any position pos The rotor wake induced airflow velocity v at pos , v pos Represented as r pos The nonlinear function is denoted by v pos =v pos (r pos ), the specific expression is: Among them, e k TPP Represents the unit vector of the blade tip trajectory plane coordinate system, k = 1, 2, 3; Yes pos Position coordinates in the blade tip trajectory plane coordinate system, j = x, y, z.

5. The method according to claim 4, characterized in that In step 1, the propeller aerodynamic model includes: The propeller flow velocity is expressed as: v pr =V+ω×r cg2pr -v mr2pr =μ pr Oh pr R pr e1 PRW +0e2 PRW +λ pr Oh pr R pr e3 PRW (12) Among them, v pr is the propeller flow velocity vector; r cg2pr represents the position vector from the center of gravity to the propeller; v mr2pr is the induced velocity of the rotor wake at the center of the propeller; e k PRW are the three unit vectors of the propeller wind axis system, k = 1, 2, 3, the XOY plane is parallel to the propeller hub, and the X axis is facing the incoming flow direction; μ pr and λ pr are the propeller advance ratio and inflow ratio respectively; Ω pr and R pr are the speed and radius of the propeller respectively; The propeller inflow velocity is obtained through momentum theory, and the equation is: Among them, C T,pr is the propeller thrust coefficient; v i,pr is the propeller induced speed; Propeller force vector F mr and the moment vector M mr Only tensile forces and reactive torques are considered.

6. The method according to claim 5, characterized in that In step 1, the wing aerodynamic model under the interference of the main rotor includes: the wing is divided into a wake area and a non-wake area; In order to calculate the wake area and non-wake area, the wing is divided into small areas using a grid. The total number of areas is N. wing , the i-th region and its center point are respectively denoted as s grid,i and r mr2wing,i , the rotor aerodynamic disturbance v in the i-th region is obtained by formula (11): mr2wing,i ; Define the set of areas affected by the rotor The total area of ​​the wing disturbed by the rotor wake is S wing,inMR and the average air velocity v in the area wing,inMR It can be expressed approximately as: Among them, r cg2wing,i represents the position vector from the center of gravity to the center point of the i-th region; Then the wing aerodynamic force F in the main rotor wake interference area is wing,inMR It is expressed as: F wing,inMR =q wing,inMR S wing,inMR C wing (16) Among them, q wing,inMR It is dynamic pressure; C wing is the aerodynamic coefficient of the wing; Get the wing aerodynamic force F outside the main rotor wake interference area wing,outMR , the total aerodynamic load on the wing is expressed as F wing,inMR With F wing,outMR sum.

7. The method according to claim 6, characterized in that In step 1, the aerodynamic model of the fuselage includes: the incoming flow velocity v of the fuselage fuslg It is expressed as: Among them, v mr2fuslg is the main rotor wake induced speed at the aerodynamic center of the fuselage; e k F Represents the unit vector of the fuselage coordinate system. The X, Y, and Z axes of the fuselage coordinate system point to the front, right, and bottom of the fuselage respectively. k = 1, 2, 3; are the velocity components corresponding to the X, Y, and Z axes, j = x, y, z; Aerodynamic force F of the fuselage fuslg and moment M fuslg It is expressed as: Among them, e k FW is the unit vector of the fuselage wind coordinate system, the X-axis of the fuselage wind coordinate system is facing the incoming flow direction, k = 1, 2, 3; q fuslg is the dynamic pressure; S fuslg and L fuslg are the equivalent area and length of the fuselage respectively; are the aerodynamic coefficients of *, *=Drag, Side, Lift, Roll, Pitch, Yaw, Drag, Side, Lift, Roll, Pitch, Yaw represent drag, lateral force, lift, rolling moment, pitching moment and yaw moment respectively.

8. The method according to claim 7, characterized in that In step 1, the tail aerodynamic model includes: considering the rotor aerodynamic interference, according to the tail aerodynamic center position r mr2tail Use formula (11) to calculate the disturbance airflow velocity v mr2tail , and then superimpose the free flow to get the tail relative flow velocity, then the tail aerodynamic force F tail It is expressed as: F tail =q tail S tail C tail (19) Among them, q tail is the dynamic pressure; S tail is the tail area; C tail is the aerodynamic coefficient.

9. The method according to claim 8, characterized in that Step 2 includes: The control surfaces of the single-rotor compound helicopter include: the main rotor collective pitch δ mr,col and longitudinal pitch control δ mr,lon , lateral pitch controlδ mr,lat 、N pr Collective pitch control of each propeller N wing Control surfaces on each wing and N tail Tail rudder control Total N tot =N pr +N wing +N tail +3 control surfaces, plus helicopter pitch and roll attitude, totaling N trim =N pr +N wing +N tail +5 unknown balance quantities; The allocation method of each control surface is established according to the channels and functions affected by different control surfaces. The obtained control allocation strategy is expressed as: Among them, δ in is the input column vector of the five manipulation channels, δ c is the control surface column vector of the single-rotor compound helicopter, [λ] is the control allocation coefficient matrix from the control surface column vector to the control channel input column vector; δ ver , δ roll , δ pitch , δ yaw , δ thrust They represent the vertical control channel, the roll control channel, the pitch control channel, the yaw control channel and the thrust control channel respectively; They represent the control distribution coefficients of the vertical, lateral and longitudinal control channels of the main rotor respectively; is the lateral control channel allocation coefficient of the wing; They represent the distribution coefficients of the propeller's yaw and thrust control channels respectively; They represent the distribution coefficients of the vertical, longitudinal and yaw control channels of the tail wing respectively; Based on the control allocation strategy, a trim calculation method is adopted to calculate the trim state of the single-rotor compound helicopter.

10. The method according to claim 9, characterized in that In step 2, the trim state of the single-rotor compound helicopter is calculated using a trim calculation method, specifically including: Step 2-1, set the initial value of the control channel input and the helicopter roll angle, input the flight speed corresponding to the trim state into the control allocation strategy and the helicopter pitch angle change curve, and obtain the control allocation coefficient matrix and the preset helicopter pitch angle at the trim speed point; Step 2-2, invert the control distribution coefficient matrix to obtain the control surface deflection corresponding to the current control channel input, input the control surface deflection and the preset helicopter pitch angle and helicopter roll angle initial values ​​into formula (1), and obtain the acceleration and angular acceleration corresponding to the initial values; Step 2-3, if the acceleration and angular acceleration are zero, the balancing is completed, otherwise the balancing amount is adjusted by the gradient descent method until the balancing is completed.

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