A general modeling and trim calculation method for single-rotor compound helicopters

Abstract

This invention provides a general modeling and trim calculation method for single-rotor compound helicopters, comprising the following steps: Step 1, establishing a model of the single-rotor compound helicopter; Step 2, solving the model established in Step 1 using the general trim calculation method for single-rotor compound helicopters. The general modeling and trim calculation method proposed in this invention employs empirical rotor wake and aerodynamic disturbance models, is independent of correction factors, and possesses sufficient accuracy. It also exhibits versatility and accuracy, and can be used for flight characteristic evaluation of single-rotor compound helicopters with different configurations during the conceptual design phase.

Description

Technical Field

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

[0002] Helicopters have been widely used in military and civilian fields due to their vertical takeoff and landing capabilities and excellent low-speed flight performance. However, the maximum flight speed of helicopters is limited by the compression effect of the advancing blade and the stall of the retreating blade of the main rotor, which in turn limits their flight envelope, application range, and mission efficiency. With the growth of civilian and military needs, there is an increasing desire to improve the cruise speed and range of helicopters without sacrificing hovering capabilities, and compound helicopters are a promising solution (Reference: R.A. 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 typically built by adding wings and auxiliary thrusters to a conventional single-rotor helicopter with a tail rotor to increase flight speed. During hovering and low-speed flight, compound helicopters behave similarly to conventional helicopters, with the main rotor providing lift and cyclic pitch control. At high speeds, the wings and auxiliary thrusters 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 maneuverability. Compared to tiltrotor aircraft, compound helicopters offer hovering and low-speed performance closer to conventional helicopters, while also being structurally simpler, having reusable production lines, and being more economical. Therefore, compound helicopters have attracted widespread global attention 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, preliminary assessment of their flight characteristics is crucial during the conceptual design phase of development, necessitating the establishment of a reliable flight dynamics model. However, the additional wings and propellers introduce complex aerodynamic interference problems, and the diverse configurations and variable aerodynamic interferences of compound helicopters make flight dynamics modeling extremely difficult and lacking in versatility. Furthermore, to ensure control effectiveness at both low and high speeds, compound helicopters possess control surfaces similar to those of helicopters and fixed-wing aircraft, resulting in redundant control. This means the number of unknowns in the trim calculation exceeds the number of equations, preventing traditional helicopter trim calculation methods from yielding a unique solution, which is also detrimental to flight characteristic assessment.

[0005] Numerous single-rotor compound helicopter configurations have been proposed, including the AH-56A, X-49A, X3, and RACER. During their development, most employ flight dynamics modeling methods based on wind tunnel test data, which are not only inefficient but also difficult to modify, limiting their applicability to a single model. Although two universal flight dynamics models suitable for single-rotor compound helicopters exist globally—the HeliUM model developed by the University of Maryland and the GenHel model developed by Sikorsky—both are difficult to obtain through public channels.

[0006] The key to flight dynamics modeling of a single-rotor compound helicopter lies in aerodynamic interference. Existing research focuses on using computational fluid dynamics (CFD) methods to study the aerodynamic interference between the main rotor, propeller, and wing, revealing the causes and effects of aerodynamic interference at the mechanistic level (References: T. Stokkermans, M. Voskuijl, L. Veldhuis, B. Soemarwoto, Aerodynamic Installation Effects of Lateral Rotors on a NovelCompound 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 RACERin Cruise Flight, J. Am. Helicopter Soc. 65 (2020) 1–14. https: / / doi.org / 10.4050 / JAHS.65.042001. However, when modeling aerodynamic disturbances, empirical methods based on correction factors are still mainly used. Although this reduces the need for wind tunnel testing, it still requires long CFD calculations. Scholars have established rotor-wing aerodynamic interference modeling methods that do not rely on correction factors, as well as downwash interference velocity modeling methods based on simple rotor wakes (References: PB Ferguson, A mathematical model for real-time flight simulation of ageneric 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 reflect the main characteristics of aerodynamic interference, but they have not yet been combined with single-rotor compound helicopter flight dynamics modeling.Furthermore, most existing trim calculation methods under redundant control are specific to a particular aircraft type (References: Y.Su, Z.Wang, Y.Cao, A hybrid trimstrategy for coaxial compound helicopter, Proc. Inst. Mech. Eng. Part GJ. Aerosp. Eng. 237 (2023) 452–466. https: / / doi.org / 10.1177 / 09544100221103021; ​​Y.Yuan, R.Chen, P.Li, Trim investigation for coaxial rigid 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), lacking a universal trim method applicable to different single-rotor compound helicopter types.

[0007] In the existing technology, most single-rotor compound helicopter flight dynamics models use wind tunnel test data or correction factors obtained from CFD calculations to model aerodynamic disturbances, resulting in high cost, low efficiency, difficulty in modification and availability. At the same time, the trim calculation method under redundant control also lacks universality, which is not conducive to the flight characteristic 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 that the existing flight dynamics models and trim calculation methods for single-rotor compound helicopters are costly and lack versatility. The invention aims to establish a universal flight dynamics model and a trim calculation method under redundant control that is suitable for flight characteristic evaluation in the conceptual design stage.

[0009] This invention provides a general modeling and trim calculation method for a single-rotor compound helicopter. Specifically, it consists of a flight dynamics modeling method composed of independent models of each aerodynamic component and a theoretical aerodynamic interference model, and a trim calculation method under redundant control pre-allocation conditions. It is applicable to the flight characteristic evaluation of single-rotor compound helicopters in the conceptual design stage.

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

[0011] Step 1: Build a model of a single-rotor compound helicopter;

[0012] Step 2: Use the general trim calculation method for single-rotor compound helicopters to solve the model established in Step 1.

[0013] Step 1 includes: The model of the single-rotor compound helicopter includes 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.

[0014] The rigid body dynamics model of the helicopter is represented as follows:

[0015]

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

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

[0018] The rotor flapping motion is described using a planar model of the blade tip trajectory. Let β(ψ,t) be the flapping angle of the blade at azimuth angle ψ and time t. It is approximated using a first-order Fourier series and time-varying coefficients, and expressed as:

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

[0020] Where a0(t) is the rotor cone angle; a1(t) and b1(t) represent the longitudinal and transverse first harmonic coefficients, respectively, and a1(t) and b1(t) correspond to the back tilt angle and left tilt angle of the blade tip trajectory plane, respectively, on a physical level.

[0021] The rotor flapping dynamics equations, expressed in terms of a0(t), a1(t), and b1(t), are as follows:

[0022]

[0023] Among them, D TPP K TPP f TPP These are the damping matrix, stiffness matrix, and right-hand side terms of the rotor flapping dynamics equations, respectively. It is the first derivative of a0, a1, and b1. It is the second derivative of a0, a1, and b1;

[0024] The main rotor inflow distribution is described using Pitt-Peters dynamic inflow theory, with azimuth angle ψ, time t, and dimensionless spanwise position denoted as follows: The vertical inflow at that point is Approximated by a 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 inflow coefficients of the first-order sinusoidal harmonic and the first-order cosine harmonic, respectively;

[0027] The rotor flapping dynamics equations use v0(t), v 1s (t), v 1c (t) is represented as:

[0028]

[0029] Where M and L are the quality matrix and the inflow gain matrix, respectively; C T C L C M These represent the aerodynamic lift, roll moment, and pitch moment of the main rotor, respectively. It is v0, v 1s v 1c The first derivative;

[0030] The force vector F of the main rotor is obtained by integrating the aerodynamic and inertial forces during the rotor rotation cycle. mr and torque 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, where the X-axis of the blade tip trajectory plane coordinate system is intersected by the free flow V. w The direction is the same as the free flow, and the angle between it and the free flow is denoted as -α. mr The Z-axis is perpendicular to the tip trajectory plane and points downwards, while the Y-axis is determined according to the right-hand rule; the evolution of the wake uses pseudo-time τ. wThis means that only the component of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane, v, is considered. iw (τ w Then, the z and x positions of the rotor wake centerline in the rotor tip trajectory plane coordinate system are defined by the derivative as:

[0032]

[0033] Where d represents the differential; x w z w These represent the x-axis and z-axis positions of the rotor wake centerline, respectively. They represent x respectively w z w The first derivative;

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

[0035] Assuming the rotor wake-induced velocity increases linearly with evolution time, the component of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane, v, is defined. iw (τ w And the z-axis position of the rotor wake centerline. w (τ w ) and x-axis position x w (τ w ) is represented as:

[0036] v iw (τ w )=A1τ w +v0 (7)

[0037]

[0038] Where A1 is v iw (τ w The derivative of the wake evolution time is obtained through the following rotor wake evolution equation:

[0039]

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

[0041] The distribution of the induced airflow velocity from the rotor wake is assumed to be uniform across a circular cross-section, and the rotor wake cross-section is assumed to remain circular during evolution, with a rotor wake radius R. w According to the continuity equation of momentum theory, we get:

[0042]

[0043] Based on formulas (7), (8), and (10), we obtain the result for any position r. pos The rotor wake induced airflow velocity v pos , will v pos Represented as r pos The nonlinear function, denoted as 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 The unit vector representing the plane coordinate system of the propeller tip trajectory, k = 1, 2, 3; is r pos The position coordinates in the plane coordinate system of the propeller tip trajectory, j = x, y, z.

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

[0053] The incoming flow velocity of the propeller 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 It is the velocity vector of the incoming flow from the propeller; r cg2pr This represents the position vector from the center of gravity to the propeller; v mr2pr It is the induced velocity of the rotor wake at the center of the propeller; e k PRW These are the three unit vectors of the propeller shaft system, k = 1, 2, 3. The XOY plane is parallel to the propeller hub, and the X-axis faces the direction of the incoming flow; μ pr and λ pr These are the propeller's forward ratio and inflow ratio, respectively; Ω pr and R pr These are the propeller's rotational speed and radius, respectively.

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

[0057]

[0058] Among them, C T,pr It is the propeller's thrust coefficient; v i,pr It is the propeller-induced velocity;

[0059] The force vector F of the propeller mr and torque vector M mr Only tensile force and counter-torque are considered.

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

[0061] To calculate the wake and non-wake regions, the wing is divided into small regions using a mesh, with a total of N regions. wing Let s be the location of the i-th region and its center point. grid,i and r mr2wing,i The rotor aerodynamic disturbance v in the i-th region is obtained through formula (11). mr2wing,i Define the set of regions affected by rotor interference. The total area S of the wing affected by the rotor wake is... wing,inMR and the average airflow velocity v in the region wing,inMR Approximately expressed as:

[0062]

[0063] Where, r cg 2wi ng, i This represents the position vector from the centroid to the center point of the i-th region;

[0064] Then, the wing aerodynamic force F within the main rotor wake interference zone wing,inMR Expressed as:

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

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

[0067] Obtain the wing aerodynamic force F outside the main rotor wake interference region. 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 airflow velocity v of the fuselage. fuslg Expressed as:

[0069]

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

[0071] aerodynamic force F of the fuselage fuslg and torque M fuslg Expressed as:

[0072]

[0073]

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

[0075] In step 1, the tail fin aerodynamic model includes: considering rotor aerodynamic interference, based on the tail fin aerodynamic center position r mr2tail The velocity of the disturbing airflow v is calculated using formula (11). mr2tail Then, by superimposing the free flow velocity to obtain the relative velocity of the tail fin to the incoming flow, the aerodynamic force F of the tail fin is obtained. tail Expressed as:

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

[0077] Where, q tail It is dynamic pressure; S tail It is the tail fin area; C tail It is the aerodynamic coefficient.

[0078] Step 2 includes: The control surfaces of the single-rotor compound helicopter include: the collective pitch δ of the main rotor. mr,col and longitudinal pitch control δ mr,lon Lateral pitch control δ mr,lat 、N pr collective pitch control of the propeller N wing Control surfaces of each wing and N tail Tail fin control surfaces 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 unknowns in the balancing process;

[0079] Based on the channels and effects of different manipulation surfaces, an allocation method for each manipulation surface is established, and the resulting manipulation allocation strategy is expressed as follows:

[0080]

[0081] Where, δ in It is the input column vector of 5 manipulation channels, δ c δ is the control surface column vector of a single-rotor compound helicopter, where [λ] is the control allocation coefficient matrix from the control surface column vector to the control channel input column vector; ver δ roll δ pitch δ yaw δ thrustThese represent the vertical control channel, roll control channel, pitch control channel, yaw control channel, and thrust control channel, respectively. These represent the control allocation coefficients for the vertical, lateral, and longitudinal control channels of the main rotor, respectively. It is the lateral control channel allocation coefficient of the wing; These represent the allocation coefficients for the propeller's yaw and thrust control channels, respectively. These represent the allocation coefficients for the vertical, longitudinal, and yaw control channels of the tail fin, respectively.

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

[0083] Step 2, which involves calculating the trim state of a single-rotor compound helicopter using a trim calculation method, specifically includes:

[0084] Step 2-1: Set the initial values ​​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 allocation coefficient matrix to obtain the control surface deflection corresponding to the current control channel input. Input the control surface deflection and the preset initial values ​​of helicopter pitch angle and helicopter roll angle into formula (1) to obtain the acceleration and angular acceleration corresponding to the initial values.

[0086] Steps 2-3: If the acceleration and angular acceleration are zero, the balancing is complete; otherwise, the balancing amount is adjusted using the gradient descent method until the balancing is complete.

[0087] Beneficial effects: Compared with existing technologies, the technical solutions adopted in this invention are mostly based on correction factors obtained from wind tunnel test data or CFD calculations. These technologies are costly, inefficient, difficult to modify, and hard to obtain. Moreover, the trim calculation methods lack universality. In contrast, the universal modeling and trim calculation method proposed in this invention uses empirical rotor wake and aerodynamic disturbance models, does not depend on correction factors, and has sufficient accuracy. It is both universal and accurate, and can be used to evaluate the flight characteristics of single-rotor compound helicopters with different configurations in the conceptual design stage. Attached Figure Description

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

[0089] Figure 1 This is a schematic diagram of a general modeling and trim calculation method for single-rotor compound helicopters.

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

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

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

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

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

[0095] This embodiment provides a general modeling and trim calculation method for single-rotor compound helicopters, applicable to the flight characteristic evaluation of single-rotor compound helicopters with different configurations during the conceptual design phase. For example... Figure 1 The diagram shows a general modeling and trim calculation method for a single-rotor compound helicopter. The upper part shows the equations of motion for the single-rotor compound helicopter, and the lower part shows the method for reducing the number of unknowns. Finally, the number of unknowns is matched with the number of equations of motion, so that a unique solution for the trim state can be obtained.

[0096] First, a general flight dynamics model for a single-rotor compound helicopter is established, consisting of a main rotor aerodynamic-flapping dynamics model, a helicopter rigid body dynamics model, a propeller aerodynamic model, a fuselage aerodynamic model, a wing aerodynamic model, a tail aerodynamic model, and an aerodynamic disturbance model. The overall equations of motion can be represented by the rigid body dynamics model.

[0097] The main rotor aerodynamic-flailing dynamics model uses a blade tip trajectory plane model to describe the rotor flapping motion and 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 and inertial forces during the rotor rotation period. mr and torque vector M mr Since all aerodynamic components of a single-rotor compound helicopter are located below the main rotor, the downwash airflow from the main rotor has a significant impact on these components during hovering and low-speed flight. This invention proposes... Figure 2 The simplified rotor wake model shown, combined with the Pitt-Peters dynamic inflow model and the blade tip trajectory plane dynamics model, can be used to obtain the rotor wake model at any position r. pos The rotor wake induced airflow velocity v posThis is used to account for 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. First, the inflow velocity of the propeller is expressed as the superposition of the propeller relative to the inflow velocity and the velocity induced by the rotor wake at the center of the propeller. Then, the inflow velocity of the propeller is obtained through momentum theory. Finally, the thrust and counter-torque of the propeller are obtained through the same calculation method as that of the main rotor.

[0099] To account for the significant impact of the main rotor wake on the wing, this invention proposes an empirical rotor-wing aerodynamic interference model without correction factors, making it applicable to modeling various single-rotor compound helicopters. Figure 3 An empirical rotor-wing aerodynamic interference model is presented. The wing is divided into a gray "wake region" and a white "non-wake region." The downwash from the main rotor primarily affects the "wake region" but has no effect on the "non-wake region." The areas of the "wake region" and "non-wake region" are calculated by dividing the wing into small regions using a grid. Thus, the wing aerodynamic forces within the main rotor wake interference region can be represented as the superposition of wing aerodynamic forces inside and outside the main rotor wake interference region.

[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 depends on whether the aerodynamic center of the fuselage is located within the rotor wake. The aerodynamic interference speed is determined by the airflow speed induced by the rotor wake at the aerodynamic center of the fuselage.

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

[0102] Then, a general trim calculation method applicable to single-rotor compound helicopters (such as...) is established. Figure 4 (As shown). The control surfaces of a single-rotor compound helicopter are classified as follows: 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 the propeller N wing Control surfaces of each wing and N tail Tail fin control surfaces 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 control allocation coefficients and pre-designing the pitch angle versus forward velocity curve, the number of trim unknowns is reduced to 6, which is equal to the 6 equations provided by the rigid body dynamics equations, thus obtaining a unique trim solution.

[0103] Finally, taking a twin-rotor propulsion compound helicopter as an example, wind tunnel trim tests were conducted on a 12kg prototype (e.g. Figure 5 (As shown). Figure 6a , Figure 6b , Figure 6c , Figure 6d , Figure 6e The model calculation results show good agreement with the experimental values, proving the rationality of the method in this embodiment.

[0104] In one specific embodiment of the present invention, a wind tunnel trim test was conducted on a 12kg prototype of a dual-propeller propulsion compound helicopter to demonstrate that the method of this embodiment can reasonably predict the flight characteristics of a single-rotor compound helicopter.

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

[0106] 1. A General Modeling Method for Single-Rotor Compound Helicopters

[0107] The single-rotor compound helicopter model consists of a rigid body dynamics model of the helicopter, an aerodynamic-flailing dynamics model of the main rotor, a main rotor wake interference model, a propeller aerodynamic model, an wing aerodynamic model under main rotor interference, a fuselage aerodynamic model, and a tail aerodynamic model. Since all aerodynamic components of the single-rotor compound helicopter are located below the main rotor, the downwash airflow from the main rotor has a significant impact on these components during hovering and low-speed flight. However, accurately considering these aerodynamic interactions usually requires expensive and time-consuming wind tunnel experiments to obtain correction factors. During the conceptual design and development phase, a generalized and simplified consideration of these interactions is more beneficial for the initial evaluation of the design scheme. The following describes a single-rotor compound helicopter flight dynamics modeling method that does not rely on correction factors.

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

[0109]

[0110] Among them, V={u,v,w} T and ω={p,q,r} T These are the translational and angular velocity vectors of the helicopter; m and I are the mass and inertia matrices of the helicopter, respectively; g is the gravity vector; F * and M * Let * represent the force vector and torque vector, respectively, where *∈{mr,fuslg,pr,wing,tail}, and mr,fuslg,pr,wing,tail represent the main rotor, fuselage, propeller, wing, and tail, respectively; n pr n wing and n tail These refer to the number of propellers, wings, and tail fins, respectively.

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

[0112] 1.1 Aerodynamic-flailing dynamics model of the main rotor

[0113] The main rotor provides the primary lift during hovering and low-speed flight. The rotor flapping motion is described using a tip-path-plane (TPP) model. Let β(ψ,t) be the flapping angle of the blade at azimuth angle ψ and time t. It can be approximated by a first-order Fourier series and time-varying coefficients, as shown below:

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

[0115] In the formula, a0(t) is the rotor cone angle; a1(t) and b1(t) represent the longitudinal and transverse first harmonic coefficients, respectively, which correspond to the backward and left tilt angles of the blade tip trajectory plane at the physical level.

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

[0117]

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

[0119] The main rotor inflow distribution is described using Pitt-Peters dynamic inflow theory, with azimuth angle ψ, time t, and dimensionless spanwise position denoted as follows: The vertical inflow at that point is It can be approximated by a first-order Fourier series and time-varying coefficients, as shown below:

[0120]

[0121] In the formula, v0(t) is the uniform inflow term of the rotor; v 1s (t) and v 1c (t) represent the inflow coefficients of the first-order sinusoidal harmonic and the first-order cosine harmonic, respectively.

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

[0123]

[0124] Where M and L are the quality matrix and the inflow gain matrix, respectively; C T C L C M These are the aerodynamic lift, roll moment, and pitch moment of the main rotor. Detailed expressions can be found in the reference: DMPitt, DAPeters, 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 and inertial forces during the rotor rotation cycle. mr and torque vector M mrFor detailed expressions, please refer to the following reference: HBHilbert, A Mathematical Model of the UH-60 Helicopter, Ames Research Center, Moffett Field, California, 1984.

[0126] 1.2 Main rotor wake interference model

[0127] This invention proposes, as follows Figure 2 The simplified rotor wake model shown is combined with the Pitt-Peters dynamic inflow model and the blade tip trajectory plane dynamics model to account for aerodynamic interference between the main rotor and other components.

[0128] The rotor wake model is defined in the rotor tip trajectory plane coordinate system, with its X-axis intersecting the free flow V. w The direction is the same as the free flow, and the angle between it and the free flow is denoted as -α. mr The Z-axis is perpendicular to the tip trajectory plane and points downwards, while the Y-axis is determined according to the right-hand rule. The evolution of the wake uses pseudo-time τ. w This 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 rotor tip trajectory plane coordinate system can be defined by their derivatives as follows:

[0129]

[0130] When τ w When v = 0, the wake-induced airflow velocity is equal to the uniform inflow magnitude v0 at the main rotor disk. According to momentum theory, the wake-induced airflow velocity increases from v0 to 2v0 at infinity. In practical applications, it can be reasonably assumed that the wake has 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 evolution time, v can be... iw (τ w ), z w (τ w ), x w (τ w The following is represented:

[0131] v iw (τ w )=A1τ w +v0 (7)

[0132]

[0133] In the formula, A1 is v iw (τ w The derivative of the wake evolution time 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 of 1.5 times the radius.

[0136] Assuming the rotor wake-induced airflow velocity distribution is uniform across a circular cross-section, and that this cross-section remains circular during evolution, its radius can be obtained from the continuity equation of momentum theory:

[0137]

[0138] Based on formulas (7), (8), and (10), we obtain the result for any position r. pos The rotor wake induced airflow velocity v pos v pos Represented as r pos The nonlinear function, denoted as 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 The unit vector representing the plane coordinate system of the propeller tip trajectory, k = 1, 2, 3; It is r pos The position coordinates in the plane coordinate system of the propeller tip trajectory, j = x, y, z.

[0147] 1.3 Propeller Aerodynamic Model

[0148] The propeller provides additional thrust to a single-rotor compound helicopter, counteracts the torque of the main rotor, and provides additional control surfaces. Its aerodynamic model is similar to that of the main rotor, but the dynamic changes in flapping and inflow are ignored. First, the inflow velocity of the propeller 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 It is the velocity vector of the incoming flow from the propeller; r cg2pr This represents the position vector from the center of gravity to the propeller; v mr2pr It is the induced velocity of the rotor wake at the center of the propeller; e k PRW These are the three unit vectors of the propeller shaft system, k = 1, 2, 3. The XOY plane is parallel to the propeller hub, and the X-axis faces the direction of the incoming flow; μ pr and λ pr These are the propeller's forward ratio and inflow ratio, respectively. Ω pr and R pr These are the propeller's rotational speed and radius, respectively.

[0151] The propeller inflow velocity is obtained using momentum theory, with the following equation:

[0152]

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

[0154] The force vector F of the propeller mr and torque vector M mr Considering only the thrust and counter-torque, the calculation method is the same as that for the main rotor.

[0155] 1.4 Aerodynamic Model of the Wing under Main Rotor Interference

[0156] During hovering and low-speed flight, the wake of the main rotor significantly affects the wing, thus requiring consideration of the aerodynamic interference between the rotor and the wing. Referring to the rotor-airfoil aerodynamic interference model for tiltrotor aircraft (references: PB Ferguson, A mathematical model for real-time flight simulation of ageneric 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), this invention proposes an empirical rotor-wing aerodynamic interference model without correction factors, making it applicable to modeling various single-rotor compound helicopters. Figure 3 An empirical rotor-wing aerodynamic interference model was presented. The wing was divided into a gray "wake region" and a white "non-wake region." The downwash airflow from the main rotor primarily affected the "wake region" but had no effect on the "non-wake region."

[0157] To calculate the "wake region" and "non-wake region," the wing is divided into small regions using a grid. The total number of regions is N. wing Let s be the location of the i-th small region and its center point. grid,i and r mr2wing,i Then the rotor aerodynamic disturbance v in the i-th region can be obtained through formula (11). mr2wing,i Define the set of regions affected by rotor interference as follows: The total area of ​​the wing affected by the rotor wake and the average airflow velocity within that area can be approximated as:

[0158]

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

[0160] The aerodynamic forces of the wing within the main rotor wake interference region can then be expressed as (ignoring the aerodynamic torque of the wing):

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

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

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

[0164] 1.5 Aerodynamic Model of the Airframe

[0165] The airframe aerodynamic model uses aerodynamic coefficient interpolation. If the fuselage aerodynamic center is located within the rotor wake, the rotor downwash velocity will be superimposed on the free airflow. The fuselage free airflow velocity v fuslg Expressed as:

[0166]

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

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

[0169]

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

[0171] 1.6 Tail fin aerodynamic model

[0172] The tail fin model is similar to the wing model, but due to its smaller area, it is not meshed when considering rotor aerodynamic interference. Instead, it is based on the aerodynamic center position r of the tail fin. mr2tail The velocity of the disturbing airflow v is calculated using formula (11). mr2tail Then, by superimposing the free flow velocity to obtain the tail fin's relative velocity to the incoming flow, the tail fin's aerodynamic force can be expressed as:

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

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

[0175] 2. General Trial Calculation Method for Single-Rotor Compound Helicopters

[0176] The control surfaces of a single-rotor compound helicopter include: the collective pitch of the main rotor (δ) mr,co l) and longitudinal and lateral pitch control (δ) mr, l on and δ mr,lat ), N pr collective pitch control of the propeller N wing Control surfaces of each wing and N tail Tail fin control surfaces 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 unknowns in the balancing process.

[0177] Since the rigid body dynamics equations can only provide six equations, to obtain a unique balancing solution, control allocation coefficients need to be introduced to reduce the number of balancing unknowns to six. Therefore, a control allocation method is established based on the channels and effects of different control surfaces. Simultaneously, a pitch angle versus forward velocity curve is pre-designed to ensure good flight characteristics. Table 1 lists the possible control allocation coefficients, all of which are functions of forward velocity.

[0178] Table 1

[0179]

[0180] Since the wing and tail control surfaces only become effective at a certain forward speed, their corresponding control allocation coefficients are set to zero during hovering, considering only the control surfaces of the main rotor and propeller. As the forward speed gradually increases, the control effectiveness of the wing and tail control surfaces gradually increases. At this point, the control allocation coefficients of the main rotor, propeller, wings, and tail are adjusted to maintain a uniform overall control effectiveness. During high-speed flight, propulsion is primarily achieved through the control surfaces of the wings and tail and the propeller; in this case, the allocation coefficient of the main rotor control surface is zero. The resulting control allocation strategy is expressed as follows:

[0181]

[0182] In the formula, δ in It is the input column vector of 5 manipulation channels, δ c is the control surface column vector of a single-rotor compound helicopter, and [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 forward flight speed.

[0183] Based on the above manipulation allocation strategy, the following is adopted: 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 inputs 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, thus a unique trim solution can be obtained.

[0184] The specific process is:

[0185] (1) Set the initial values ​​of the control channel input and the helicopter roll angle, input the flight speed corresponding to the trim state to 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.

[0186] (2) Invert the control allocation coefficient matrix to obtain the control surface deflection corresponding to the current control channel input. Input it and the preset initial values ​​of helicopter pitch angle and helicopter roll angle 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 complete; otherwise, the balancing amount is adjusted by the gradient descent method until the balancing is complete.

[0188] 3. Trial results of a sample single-rotor compound helicopter

[0189] To verify the correctness of the general modeling and trimming method for the single-rotor compound helicopter described above, wind tunnel trimming test data from a 12kg twin-propeller-driven compound helicopter prototype were used to validate the trimming results (e.g., Figure 5 As shown in the figure, Table 2 shows the prototype parameters.

[0190] Table 2

[0191]

[0192] The wind tunnel trim test was conducted in the low-speed wind tunnel of the National Key Laboratory of Helicopter Dynamics at Nanjing University of Aeronautics and Astronautics, with relative incoming flow velocities of 0 m / s, 5 m / s, 10 m / s, 15 m / s, and 20 m / s, and a pitch angle of 2°. During the test, the control distribution coefficients of the wing and tail 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. The control values ​​of each control surface were recorded at this point. Figure 6a , Figure 6b , Figure 6c , Figure 6d , Figure 6e The comparison between the model calculation results and the experimental values ​​shows that they are in good agreement, indicating that the general modeling and trimming method for single-rotor compound helicopters established by the method in this embodiment has good accuracy and can be used for preliminary flight characteristic assessment.

[0193] This invention provides a general modeling and trim calculation method for a single-rotor compound helicopter. Many methods and approaches exist for implementing this technical solution; the above description is merely a preferred embodiment of the invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of this invention, and these improvements and modifications should also be considered within the scope of protection of this invention. All components not explicitly stated in this embodiment can be implemented using existing technologies.

Claims

1. A general modeling and trim calculation method for a single-rotor compound helicopter, characterized in that, Includes the following steps: Step 1: Build a model of a single-rotor compound helicopter; Step 2: Use the general trim calculation method for single-rotor compound helicopters to solve the model established in Step 1; Step 1 includes: The model of the single-rotor compound helicopter includes 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. The rigid body dynamics model of the helicopter is represented as follows: Among them, V={u,v,w} T and ω={p,q,r} T These are the translational velocity and angular velocity vectors of the helicopter, respectively. u, v, w are the translational velocities along the x-axis, y-axis, and z-axis of the helicopter's body axis, respectively, and p, q, r are the angular velocities along the x-axis, y-axis, and z-axis of the helicopter's body axis, respectively. and Let f(x) represent the derivatives of the helicopter's translational velocity vector and angular velocity vector, respectively; m and I are the helicopter's mass and inertia matrices, respectively; g is the gravity vector; F * and M * Let * represent the force vector and torque vector, respectively, where *∈{mr,fuslg,pr,wing,tail}, and mr,fuslg,pr,wing,tail represent the main rotor, fuselage, propeller, wing, and tail, respectively; n pr n wing and n tail These refer to the number of propellers, the number of wings, and the number of tail fins; In step 1, the main rotor aerodynamic-flailing dynamics model includes: The rotor flapping motion is described using a planar model of the blade tip trajectory. Let β(ψ,t) be the flapping angle of the blade at azimuth angle ψ and time t. It is approximated using a first-order Fourier series and time-varying coefficients, and expressed as: β(ψ,t)=a0(t)-a1(t)cosψ-b1(t)sinψ (2) Where a0(t) is the rotor cone angle; a1(t) and b1(t) represent the longitudinal and transverse first harmonic coefficients, respectively, and a1(t) and b1(t) correspond to the back tilt angle and left tilt angle of the blade tip trajectory plane, respectively, on a physical level. The rotor flapping dynamics equations, expressed in terms of a0(t), a1(t), and b1(t), are as follows: Among them, D TPP K TPP f TPP These are the damping matrix, stiffness matrix, and right-hand side terms of the rotor flapping dynamics equations, respectively. It is the first derivative of a0, a1, and b1. It is the second derivative of a0, a1, and b1; The Pitt-Peters dynamic inflow theory is used to describe the main rotor inflow distribution. Let the vertical inflow at the azimuth angle ψ, time t, and dimensionless spanwise position r be denoted as... Approximated by a 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 inflow coefficients of the first-order sinusoidal harmonic and the first-order cosine harmonic, respectively; The rotor flapping dynamics equations use v0(t), v 1s (t), v 1c (t) is represented as: Where M and L are the quality matrix and the inflow gain matrix, respectively; C T C L C M These represent the aerodynamic lift, roll moment, and pitch moment of the main rotor, respectively. It is v0, v 1s v 1c The first derivative; The force vector F of the main rotor is obtained by integrating the aerodynamic and inertial forces during the rotor rotation cycle. mr and torque vector M mr ; In step 1, the main rotor wake interference model includes: the rotor wake model is defined in the blade tip trajectory plane coordinate system, where the X-axis of the blade tip trajectory plane coordinate system is intersected by the free flow V. w The direction is the same as the free flow, and the angle between it and the free flow is denoted as -α. mr The Z-axis is perpendicular to the tip trajectory plane and points downwards, while the Y-axis is determined according to the right-hand rule; the evolution of the wake uses pseudo-time τ. w This means that only the component of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane, v, is considered. iw (τ w Then, the z and x positions of the rotor wake centerline in the rotor tip trajectory plane coordinate system are defined by the derivative as: Where d represents the differential; x w z w These represent the x-axis and z-axis positions of the rotor wake centerline, respectively. They represent x respectively w z w The first derivative; When τ w When = 0, the wake-induced airflow velocity is equal to the uniform inflow magnitude v0 at the main rotor disk; Assuming the rotor wake-induced velocity increases linearly with evolution time, the component of the wake-induced airflow velocity perpendicular to the blade tip trajectory plane, v, is defined. iw (τ w And the z-axis position of the rotor wake centerline. w (τ w ) and x-axis position x w (τ w ) is represented as: v iw (t w )=A1τ w +v0(7) Where A1 is v iw (τ w The derivative of the wake evolution time is obtained through the following rotor wake evolution equation: Where, τ w,far It is the evolution time corresponding to the rotor wake evolving to a distance of 1.5 times the radius; R mr It is the main rotor radius; The distribution of the induced airflow velocity from the rotor wake is assumed to be uniform across a circular cross-section, and the rotor wake cross-section is assumed to remain circular during evolution, with a rotor wake radius R. w According to the continuity equation of momentum theory, we get: Based on formulas (7), (8), and (10), we obtain the result for any position r. pos The rotor wake induced airflow velocity v pos , will v pos Represented as r pos The nonlinear function, denoted as v pos =v pos (r pos The specific expression is: v pos =v pos (r pos ) Given but like Then (11) v pos =(A1τ w,pos +v0)e3 TPP otherwise v pos ＝0 Among them, e k TPP The unit vector representing the plane coordinate system of the propeller tip trajectory, k = 1, 2, 3; It is r pos The position coordinates in the plane coordinate system of the propeller tip trajectory, j = x, y, z; In step 1, the propeller aerodynamic model includes: The incoming flow velocity of the propeller 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 It is the velocity vector of the incoming flow from the propeller; r cg2pr This represents the position vector from the center of gravity to the propeller; v mr2pr It is the induced velocity of the rotor wake at the center of the propeller; e k PRW These are the three unit vectors of the propeller shaft system, k = 1, 2, 3. The XOY plane is parallel to the propeller hub, and the X-axis faces the direction of the incoming flow; μ pr and λ pr These are the propeller's forward ratio and inflow ratio, respectively; Ω pr and R pr These are the propeller's rotational speed and radius, respectively. The propeller inflow velocity is obtained through momentum theory, and the equation is: Among them, C T,pr It is the propeller's thrust coefficient; v i,pr It is the propeller-induced velocity; The force vector F of the propeller mr and torque vector M mr Considering only tension and counter-torque; In step 1, the aerodynamic model of the wing under main rotor interference includes: the wing is divided into a wake region and a non-wake region; To calculate the wake and non-wake regions, the wing is divided into small regions using a mesh, with a total of N regions. wing Let s be the location of the i-th region and its center point. grid,i and r mr2wing,i The rotor aerodynamic disturbance v in the i-th region is obtained through formula (11). mr2wing,i Define the set of regions affected by rotor interference. The total area S of the wing affected by the rotor wake is... wing,inMR and the average airflow velocity v in the region wing,inMR Approximately expressed as: Where, r cg2wing,i This represents the position vector from the centroid to the center point of the i-th region; Then, the wing aerodynamic force F within the main rotor wake interference zone wing,inMR Represented as: F wing,inMR =q wing,inMR S wing,inMR C wing (16) Where, q wing,inMR It is dynamic pressure; C wing It is the aerodynamic coefficient of the wing; Obtain the wing aerodynamic force F outside the main rotor wake interference region. wing,outMR The total aerodynamic load of the wing is expressed as F. wing,inMR With F wing,outMR sum.

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

3. The method according to claim 2, characterized in that, In step 1, the tail fin aerodynamic model includes: considering rotor aerodynamic interference, based on the tail fin aerodynamic center position r mr2tail The velocity of the disturbing airflow v is calculated using formula (11). mr2tail Then, by superimposing the free flow velocity to obtain the relative velocity of the tail fin to the incoming flow, the aerodynamic force F of the tail fin is obtained. tail Represented as: F tail =q tail S tail C tail (19) Where, q tail It is dynamic pressure; S tail It is the tail fin area; C tail It is the aerodynamic coefficient.

4. The method according to claim 3, characterized in that, Step 2 includes: The control surfaces of the single-rotor compound helicopter include: the collective pitch δ of the main rotor. mr,col and longitudinal pitch control δ mr,lon Lateral pitch control δ mr,lat N pr collective pitch control of the propeller N wing Control surfaces of each wing and N tail Tail fin control surfaces 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 unknowns in the balancing process; Based on the channels and effects of different manipulation surfaces, an allocation method for each manipulation surface is established, and the resulting manipulation allocation strategy is expressed as follows: Where, δ in It is the input column vector of 5 manipulation channels, δ c δ is the control surface column vector of a single-rotor compound helicopter, where [λ] is the control allocation coefficient matrix from the control surface column vector to the control channel input column vector; ver δ roll δ pitch δ yaw δ thrust These represent the vertical control channel, roll control channel, pitch control channel, yaw control channel, and thrust control channel, respectively. These represent the control allocation coefficients for the vertical, lateral, and longitudinal control channels of the main rotor, respectively. It is the lateral control channel allocation coefficient of the wing; These represent the allocation coefficients for the propeller's yaw and thrust control channels, respectively. These represent the allocation coefficients for the vertical, longitudinal, and yaw control channels of the tail fin, respectively. Based on the aforementioned control allocation strategy, the trim calculation method is used to calculate the trim state of the single-rotor compound helicopter.

5. The method according to claim 4, characterized in that, Step 2, which involves calculating the trim state of a single-rotor compound helicopter using a trim calculation method, specifically includes: Step 2-1: Set the initial values ​​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 allocation coefficient matrix to obtain the control surface deflection corresponding to the current control channel input. Input the control surface deflection and the preset initial values ​​of helicopter pitch angle and helicopter roll angle into formula (1) to obtain the acceleration and angular acceleration corresponding to the initial values. Steps 2-3: If the acceleration and angular acceleration are zero, the balancing is complete; otherwise, the balancing amount is adjusted using the gradient descent method until the balancing is complete.

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