A landing flight characteristics analysis method for coaxial rigid rotor helicopters

Abstract

The present invention discloses a landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter, comprising: step 1, using a high-precision detached vortex method to obtain high-confidence unsteady stern flow field data; step 2, based on the unsteady stern flow field data, constructing a flight dynamics model of the coaxial rigid rotor helicopter for coupling the unsteady stern flow field data; step 3, coupling the unsteady stern flow field data to the flight dynamics model of the coaxial rigid rotor helicopter, and calculating the control margin and pilot workload of the coaxial rigid rotor helicopter. The technical solution provided by the present invention solves the problem that there is a gap in the research on the landing flight characteristics of the coaxial rigid rotor helicopter in the existing analysis scheme of the landing flight characteristics of the helicopter, and the problem that the analysis accuracy of the landing flight characteristics of the existing conventional coaxial helicopter is low.

Description

Technical Field

[0001] The present invention relates to the field of, but is not limited to, helicopter landing flight mechanics technology, and in particular to a landing flight characteristics analysis method suitable for a coaxial rigid rotor helicopter. Background Art

[0002] At present, the research in the field of helicopter landing mainly focuses on conventional single-rotor helicopters. In the single-rotor helicopter landing technology, the ship flow field database is first obtained through empirical models, experiments or numerical simulations; then the ship flow field database is coupled to the helicopter flight dynamics model through data transfer strategies, so as to analyze the landing flight characteristics.

[0003] In the analysis of coaxial twin-rotor helicopter landing technology, only Yang Jun in China has conducted a preliminary study on the balance characteristics of conventional coaxial helicopters under the influence of steady stern flow field in the currently published literature (Yang Jun. Calculation of wind limit diagrams for coaxial helicopter takeoff and landing on ship deck. Journal of Naval Aeronautical Engineering College, 2017, 32(01):149-153.). In his study, the steady flow field data in the ship landing area was first obtained through CFD calculation, and then the steady flow field data was directly added to the rotor induced speed to obtain the rotor aerodynamic force, thereby conducting a coaxial helicopter balance characteristic analysis.

[0004] From the above analysis of helicopter landing flight characteristics, it can be seen that the current research programs of helicopter landing analysis technology at home and abroad only cover the flight characteristics analysis of single-rotor helicopter landing and conventional coaxial helicopter landing, and there is a gap in the research of coaxial rigid rotor helicopter landing flight characteristics; in addition, the analysis accuracy of existing conventional coaxial helicopter landing flight characteristics is relatively low. Summary of the invention

[0005] The purpose of the present invention is: the embodiment of the present invention provides a landing flight characteristics analysis method suitable for a coaxial rigid rotor helicopter, so as to solve the problem that there is a gap in the research on the landing flight characteristics of coaxial rigid rotor helicopters in the existing analysis schemes of helicopter landing flight characteristics, and the analysis accuracy of the landing flight characteristics of existing conventional coaxial helicopters is low.

[0006] The technical solution of the present invention is:

[0007] An embodiment of the present invention provides a method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, comprising:

[0008] Step 1, using a high-precision detached vortex method to obtain high-confidence unsteady stern flow field data;

[0009] Step 2, based on the unsteady stern flow field data, construct a flight dynamics model of a coaxial rigid rotor helicopter for coupling the unsteady stern flow field data;

[0010] Step 3, coupling the unsteady stern flow field data to the flight dynamics model of the coaxial rigid rotor helicopter, and calculating the control margin and pilot workload of the coaxial rigid rotor helicopter.

[0011] Optionally, in the above-mentioned method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, step 1 comprises:

[0012] Step 11, generating a rectangular computational domain grid with a preset length, width, and height according to the shape and size of the ship;

[0013] Step 12, setting the inlet boundary of the rectangular computational domain grid as the velocity inlet condition, and setting the outlet boundary as the pressure outlet condition;

[0014] Step 13, setting the hull as a non-slip wall, and setting the sea level and other computational domain surfaces as sliding walls;

[0015] Step 14, setting the thickness of the first layer of the boundary layer grid on the surface of the ship to 1 mm to 2 mm, and setting the boundary layer grid growth rate to 1.05 to 1.15, so as to obtain the total number of grids of the ship calculation model;

[0016] Step 15, according to the calculation and setting values ​​of steps 11 to 14, a high-precision detached vortex method is used to calculate the unsteady stern flow field data of a preset time to obtain a stable unsteady stern flow field topology structure;

[0017] Step 16, after obtaining the stable unsteady stern flow field topology structure, calculate the unsteady stern flow field data at a specified time, and output the 7-dimensional unsteady stern flow field data.

[0018] Optionally, in the above-mentioned landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter, the coaxial rigid rotor helicopter adopts a layout in which the upper rotor and the lower rotor are coaxial and reversed; the step 2 comprises:

[0019] Step 21, respectively determining the interference factors of the upper rotor and the lower rotor, so as to obtain and quantitatively calculate the mutual interference intensity between the upper rotor and the lower rotor;

[0020] Step 22, determining the inflow models of the upper rotor and the lower rotor respectively according to the interference factor, so as to represent the induced velocity distribution of the upper rotor and the lower rotor;

[0021] Step 23, determining the aerodynamic loads of the upper rotor and the lower rotor respectively;

[0022] Step 24, respectively determining the flapping motion models of the upper rotor and the lower rotor;

[0023] Step 25, respectively determining aerodynamic models of the fuselage, horizontal tail, vertical tail and tail rotor, wherein the aerodynamic models include aerodynamic forces and moments;

[0024] Step 26, determining the flight dynamics model of the coaxial rigid rotor helicopter.

[0025] Optionally, in the above-mentioned method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, step 21 comprises:

[0026] The interference factors of the upper rotor and the lower rotor are deduced by the slipstream theory:

[0027] δ1=S′1 / S f(d,χ2);

[0028] δ2=S′2 / S f(-d,χ1);

[0029] Where S is the rotor disc area, S′1 is the interference area between the upper rotor wake and the lower rotor disc, S′2 is the interference area between the lower rotor induced flow and the upper rotor disc, V x is the forward flight speed of the helicopter, d is the distance between the upper rotor and the lower rotor, χ1 and χ2 are the trail inclination angles of the lower rotor and the upper rotor respectively;

[0030] The step 22 comprises:

[0031] According to the interference factor in step 21, the inflow models of the upper rotor and the lower rotor are determined as follows:

[0032]

[0033]

[0034] Among them, r is the distance from the blade micro-segment to the center of the hub, R is the blade radius, K1 and K2 are the induced velocity distortion factors, ψ1 and ψ2 are the phase angles of the lower rotor and upper rotor blades respectively.

[0035] Optionally, in the above-mentioned method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, step 23 comprises:

[0036] A single blade is divided into multiple micro-segments, and unsteady stern flow field data is imported into the center of each micro-segment to obtain the tangential velocity component, spanwise velocity component and normal velocity component of each micro-segment center, and each velocity component is used to calculate the corresponding micro-segment aerodynamic force, and the aerodynamic force of each micro-segment is accumulated to obtain the aerodynamic force and moment of the root of the single blade; the aerodynamic force and moment of each blade in the upper rotor or the lower rotor are summed respectively to obtain the aerodynamic load of the upper rotor or the lower rotor.

[0037] Optionally, in the above-mentioned method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, step 24 comprises:

[0038] a. The first-order rigid flapping motion of the equivalent articulated rotor is used to fit the first-order elastic flapping motion of the coaxial rigid rotor; the equivalent flapping hinge offset is the intersection of the outer end tangent of the first-order elastic flapping vibration mode and the x-axis (span direction of the blade), and the stiffness of the constrained torsion spring applied at the equivalent flapping hinge is required to be:

[0039]

[0040] Where, e is the equivalent flapping offset, M β ,I β are the blade mass static moment and inertia moment respectively, Ω is the rotor speed, ω n is the first-order flapping frequency of the coaxial rigid rotor;

[0041] b, according to the centrifugal moment M CF , swinging inertia moment M I , aerodynamic moment M A 、Coriolis moment M caused by the angular velocity of the helicopter Cor 、Inertia moment M caused by fuselage angular acceleration BA , M caused by the body acceleration BL , blade gravity moment M R , and the torque M generated by the torsion spring R , the algebraic sum at the equivalent flapping hinge should be zero, and a flapping motion model of the coaxial rigid rotor is established.

[0042] Optionally, in the above-mentioned method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter, step 26 comprises:

[0043] a. The aerodynamic forces and moments of all components are synthesized at the center of gravity of the fuselage to obtain the resultant force and moment acting on the center of gravity of the helicopter in the fuselage coordinate system, and the rigid body dynamics model of the movement of the center of gravity of the helicopter and its rotation around the center of gravity is obtained;

[0044] b. Based on the flapping motion model and rigid body dynamics model of the coaxial rigid rotor, the flight dynamics model of the coaxial rigid rotor helicopter is obtained.

[0045] Optionally, in the landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter as described above, each micro-segment of each blade of the upper rotor and the lower rotor, as well as the fuselage, the horizontal tail, the vertical tail and the tail rotor are respectively an aerodynamic load calculation point, and in step 3, the interference of the unsteady stern flow field on the helicopter is simulated by importing unsteady stern flow field data into each aerodynamic load calculation point; step 3 specifically includes:

[0046] Step 31: Divide one cycle of the blade movement into N P The station is set at the station corresponding to time t0, and the instantaneous velocity field of the unsteady stern flow field is W gust (t0,u G0 ,v G0 ,w G0 );

[0047] Step 32: The coordinates of one of the micro-segments of the blade in the blade coordinate system are (t0, x b0 ,y b0 ,z b0 ), then the coordinates of the blade micro-segment in the body coordinate system are:

[0048]

[0049] The coordinates of the blade micro-segment converted into the ship coordinate system are:

[0050]

[0051] Among them, T Rb represents the transformation matrix from the blade coordinate system to the rotor rotating coordinate system, T HR represents the transformation matrix from the rotor rotating coordinate system to the hub coordinate system, T BH represents the transformation matrix from the hub coordinate system to the body coordinate system, T EB represents the transformation matrix from the body coordinate system to the inertial coordinate system, T CE represents the transformation matrix from the inertial coordinate system to the ship coordinate system; (x MR ,y MR ,z MR ) is the coordinate of the rotor hub center;

[0052] Step 33, interpolate the unsteady stern flow field data obtained in step 1 using the distance derivative weighted method to obtain the disturbance velocity component (t0, u GC0 ,v GC0 ,w GC0 ), and transform the velocity component to the blade coordinate system. After the transformation, the instantaneous velocity component of the micro-segment center is:

[0053]

[0054] Among them, T bR , T RH , T HB , T BE and T EC They are the inverse matrices corresponding to the above coordinate transformation matrices respectively;

[0055] Step 34, repeating steps 32 to 33 to obtain instantaneous velocity components of the centers of other micro-segments on the blade;

[0056] Step 35, the blade phase angle is advanced forward by one station, and the time remains unchanged, and steps 32 to 34 are repeated until the blade completes one rotational motion, and the instantaneous velocity component of each micro-segment in the blade at each station is obtained; and each instantaneous velocity component obtained is calculated and averaged respectively, so as to obtain the average unsteady stern flow field velocity component in the propeller disk plane at this moment:

[0057]

[0058] in, It represents the instantaneous velocity component obtained by interpolation of the Ith micro-segment on the blade at the K station; the average unsteady stern flow field velocity component is used to calculate the control margin of the coaxial rigid rotor helicopter.

[0059] Optionally, in the landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter as described above, step 3 further comprises:

[0060] Step 36, according to the calculation results of steps 34 and 35, the aerodynamic force and moment of the blade at time t0 are obtained, and steps 32 to 34 are repeated to obtain the aerodynamic force and moment of each blade, and the aerodynamic force and moment of all blades are summed respectively to obtain the aerodynamic force and moment of the entire rotor at time t0;

[0061] Step 37, the blade phase angle is advanced forward by one station and the time is advanced forward by Δt, and the unsteady stern flow field data of the blade micro-segment at the time t0+Δt is interpolated to obtain the instantaneous velocity component of the center of the blade micro-segment at the time t0+Δt;

[0062] Step 38, by repeating steps 32 to 34, and steps 36 to 37, the instantaneous velocity components of each micro-segment in each blade at each position and at each moment are obtained, thereby simulating the unsteady interference of the coaxial rigid rotor with the turbulent vortex structure during the periodic motion.

[0063] The beneficial effects of the present invention are:

[0064] The present invention proposes a landing flight characteristics analysis method suitable for coaxial rigid rotor helicopters, and its core content mainly includes three parts: high-precision unsteady stern flow field data acquisition, establishment of coaxial rigid rotor helicopter flight dynamics modeling that can couple unsteady stern flow field data, and data transmission strategy from unsteady stern flow field data to the flight dynamics model. In the technical solution of the embodiment of the present invention, the high-precision unsteady stern flow field obtained in the first part can be imported into the coaxial rigid rotor helicopter flight dynamics model established in the second part; the "data transmission strategy based on unidirectional coupling" in the third part is the key innovation of the present invention. In this part of the technology, the present invention adopts the idea of ​​"nested grids" to discretize the blade rotation motion into N P At each station, the unsteady interference between the coaxial rigid rotor and the large-scale vortex of the stern flow is achieved by matching the blade motion with the unsteady stern flow field on a time scale.

[0065] The landing flight characteristics analysis method for a coaxial rigid rotor helicopter provided in the embodiment of the present invention can quickly and accurately provide the changing characteristics of the control margin and the pilot workload during the landing of the coaxial rigid rotor helicopter. In addition, the simulation cycle for a landing process in the embodiment of the present invention is only about 10 days. Compared with the sea test flight, the present invention can significantly shorten the test cycle, save manpower and material resources, and reduce risks. BRIEF DESCRIPTION OF THE DRAWINGS

[0066] The accompanying drawings are used to provide a further understanding of the technical solution of the present invention and constitute a part of the specification. Together with the embodiments of the present application, they are used to explain the technical solution of the present invention and do not constitute a limitation on the technical solution of the present invention.

[0067] Figure 1 A flow chart of a method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter provided in an embodiment of the present invention;

[0068] FIG2 is a schematic diagram of the mutual interference between the upper rotor and the lower rotor in a coaxial rigid rotor helicopter based on the slipstream theory in an embodiment of the present invention;

[0069] Figure 3 Schematic diagram of the micro-segment division of a blade in an embodiment of the present invention;

[0070] Figure 4 Schematic diagram of the principle of a coaxial rigid rotor and its equivalent flapping hinge in an embodiment of the present invention;

[0071] Figure 5 It is a schematic diagram of the rotor and the fuselage being discretized into a number of calculation points of aerodynamic loads in an embodiment of the present invention;

[0072] Figure 6It is a schematic diagram of the stern flow field data domain established for the lateral shifting process in an embodiment of the present invention;

[0073] Figure 7 It is a schematic diagram of the variation curves of the control amount, attitude angle and required power when the coaxial rigid rotor helicopter performs lateral displacement at different altitudes in an embodiment of the present invention;

[0074] Figure 8 Schematic diagram of a curve showing a change in the unsteady load level of a coaxial rigid rotor helicopter at different lateral displacement path heights in an embodiment of the present invention. DETAILED DESCRIPTION

[0075] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those skilled in the art, without departing from the concept of the present invention, several variations and improvements may be made. These all belong to the protection scope of the present invention.

[0076] It has been explained in the above background technology that the existing research schemes for the analysis of the flight characteristics of helicopter landing only cover the flight characteristics analysis of single-rotor helicopter landing and conventional coaxial helicopter landing, and there is a gap in the research on the landing flight characteristics of coaxial rigid rotor helicopters; in addition, the analysis accuracy of the existing conventional coaxial helicopter landing flight characteristics is low.

[0077] Specifically, the current research on helicopter landing analysis technology at home and abroad mainly has the following deficiencies:

[0078] (1) The landing analysis technology of conventional single-rotor helicopters is relatively mature, but there is little research on the landing analysis technology of coaxial helicopters, and there is even less research on the landing flight characteristics of coaxial rigid rotor helicopters;

[0079] (2) The accuracy of the analysis of the landing flight characteristics of conventional coaxial helicopters is not enough: on the one hand, the Reynolds-averaged NS equations were used to solve the ship flow field in their study, which could not obtain high-precision unsteady flow field data; on the other hand, in their study, the ship flow field data line shape was directly added to the induced velocity, which could not simulate the interference characteristics between the blades and the ship's turbulent vortex structure.

[0080] In order to solve the above-mentioned deficiencies in the existing analysis of helicopter landing flight characteristics, an embodiment of the present invention provides a landing flight characteristics analysis method suitable for a coaxial rigid rotor helicopter. The landing flight characteristics analysis method adopts a high-precision separated vortex method to obtain high-confidence unsteady stern flow field data, and adopts a discrete blade method to couple the unsteady stern flow field data to the aerodynamic center of the coaxial rigid rotor helicopter, thereby more realistically simulating the unsteady interference of the stern flow on the helicopter.

[0081] The present invention provides the following specific embodiments which can be combined with each other. The same or similar concepts or processes may not be described in detail in some embodiments.

[0082] Figure 1 A flowchart of a method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter provided in an embodiment of the present invention is provided. The method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter provided in an embodiment of the present invention may include the following steps:

[0083] Step 1, using a high-precision detached vortex method to obtain high-confidence unsteady stern flow field data;

[0084] Step 2, based on the unsteady stern flow field data, construct a flight dynamics model of a coaxial rigid rotor helicopter for coupling the unsteady stern flow field data;

[0085] Step 3, coupling the unsteady stern flow field data to the flight dynamics model of the coaxial rigid rotor helicopter, and calculating the control margin and pilot workload of the coaxial rigid rotor helicopter.

[0086] In the embodiment of the present invention, the specific implementation process of the above step 1 may include the following steps:

[0087] Step 11, generating a rectangular computational domain grid with a preset length, width, and height according to the shape and size of the ship;

[0088] Step 12, setting the inlet boundary of the rectangular computational domain grid as the velocity inlet condition, and setting the outlet boundary as the pressure outlet condition;

[0089] Step 13, setting the hull as a non-slip wall, and setting the sea level and other computational domain surfaces as sliding walls;

[0090] Step 14, setting the thickness of the first layer of the boundary layer grid on the surface of the ship to 1 mm to 2 mm, and setting the boundary layer grid growth rate to 1.05 to 1.15, so as to obtain the grid size around the ship and the total number of grids of the ship calculation model;

[0091] Step 15, according to the calculation and setting values ​​of steps 11 to 14, a high-precision detached vortex method is used to calculate the unsteady stern flow field data for a preset time (e.g., 5s to 10s) to obtain a stable unsteady stern flow field topology structure;

[0092] Step 16, after obtaining the stable unsteady stern flow field topology structure, calculate the unsteady stern flow field data for a specified time (for example, 30 seconds), and output the 7-dimensional unsteady stern flow field data.

[0093] It should be noted that the coaxial rigid rotor helicopter in the embodiment of the present invention adopts a layout in which the upper rotor and the lower rotor are coaxial and reversed; based on the rotor structure of the coaxial rigid rotor helicopter, the specific implementation process of step 2 in the embodiment of the present invention may include the following steps:

[0094] Step 21, respectively determining the interference factors of the upper rotor and the lower rotor, so as to obtain and quantitatively calculate the mutual interference intensity between the upper rotor and the lower rotor;

[0095] Step 22, determining the inflow models of the upper rotor and the lower rotor respectively according to the interference factor, so as to represent the induced velocity distribution of the upper rotor and the lower rotor;

[0096] Step 23, determine the aerodynamic loads of the upper rotor and the lower rotor respectively,

[0097] Step 24, respectively determining the flapping motion models of the upper rotor and the lower rotor, includes:

[0098] Step 25, respectively determining aerodynamic models of the fuselage, horizontal tail, vertical tail and tail rotor, wherein the aerodynamic models include aerodynamic forces and moments;

[0099] Step 26, determining the flight dynamics model of the coaxial rigid rotor helicopter.

[0100] In one implementation of the embodiment of the present invention, as shown in FIG2 , it is a schematic diagram of the mutual interference between the upper rotor and the lower rotor in the coaxial rigid rotor helicopter based on the slipstream theory in the embodiment of the present invention. Figure 2a Specifically shows the interference of the upper rotor on the lower rotor. Figure 2b The interference of the lower rotor to the upper rotor is specifically shown. The specific implementation of the above step 21 may include:

[0101] The interference factors of the upper rotor and the lower rotor are deduced by the slipstream theory:

[0102] δ1=S′1 / S f(d,χ2);

[0103] δ2=S′2 / S f(-d,χ1);

[0104] Where S is the rotor disc area, S′1 is the interference area between the upper rotor wake and the lower rotor disc, S′2 is the interference area between the lower rotor induced flow and the upper rotor disc, V x is the forward flight speed of the helicopter, d is the distance between the upper rotor and the lower rotor, χ1 and χ2 are the wake inclination angles of the lower rotor and the upper rotor respectively.

[0105] Accordingly, in this implementation, step 22 includes:

[0106] According to the interference factor in step 21, the inflow models of the upper rotor and the lower rotor are determined as follows:

[0107]

[0108]

[0109] Among them, r is the distance from the blade micro-segment to the center of the hub, R is the blade radius, K1 and K2 are the induced velocity distortion factors, ψ1 and ψ2 are the phase angles of the lower rotor and upper rotor blades respectively.

[0110] In one implementation of the embodiment of the present invention, the implementation of step 23 includes: dividing a single blade into a plurality of micro-segments, such as Figure 3 As shown, it is a schematic diagram of the micro-segment division of the blade in the embodiment of the present invention. The unsteady stern flow field data is imported into the center of each micro-segment to obtain the tangential velocity component, the spanwise velocity component and the normal velocity component of the center of each micro-segment. The corresponding micro-segment aerodynamic force is calculated by using each velocity component. The aerodynamic force of each micro-segment is accumulated to obtain the aerodynamic force and moment of the root of a single blade. The aerodynamic force and moment of each blade in the upper rotor or the lower rotor are summed respectively to obtain the aerodynamic load of the upper rotor or the lower rotor respectively.

[0111] In an implementation of the embodiment of the present invention, the implementation process of the above step 24 may include:

[0112] a, the first-order rigid flapping motion of the equivalent articulated rotor is used to fit the first-order elastic flapping motion of the coaxial rigid rotor; where the equivalent flapping hinge offset is the intersection of the outer end tangent of the first-order elastic flapping vibration mode and the x-axis (span direction of the blade), such as Figure 4 , which is a schematic diagram of the principle of the coaxial rigid rotor and its equivalent flapping hinge in an embodiment of the present invention. The stiffness of the constrained torsion spring applied at the equivalent flapping hinge is specifically:

[0113]

[0114] Where, e is the equivalent flapping offset, M β ,I β are the blade mass static moment and inertia moment respectively, Ω is the rotor speed, ω n is the first-order flapping frequency of the coaxial rigid rotor;

[0115] b, according to the centrifugal moment M CF , swinging inertia moment M I , aerodynamic moment M A 、Coriolis moment M caused by the angular velocity of the helicopter Cor 、Inertia moment M caused by fuselage angular acceleration BA , M caused by the body acceleration BL , blade gravity moment MR , and the torque M generated by the torsion spring R , the algebraic sum at the equivalent flapping hinge should be zero, and a flapping motion model of the coaxial rigid rotor is established.

[0116] In one implementation of the embodiment of the present invention, the implementation process of the above step 26 may include:

[0117] a. The aerodynamic forces and moments of all components are synthesized at the center of gravity of the fuselage to obtain the resultant force and moment acting on the center of gravity of the helicopter in the fuselage coordinate system, and the rigid body dynamics model of the movement of the center of gravity of the helicopter and its rotation around the center of gravity is obtained;

[0118] b. Based on the flapping motion model and rigid body dynamics model of the coaxial rigid rotor, the flight dynamics model of the coaxial rigid rotor helicopter is obtained.

[0119] In the embodiment of the present invention, each micro-segment of each blade of the upper rotor and the lower rotor of the coaxial rigid rotor helicopter, as well as the fuselage, the horizontal tail, the vertical tail and the tail rotor are respectively an aerodynamic load calculation point, such as Figure 5 As shown, it is a schematic diagram of the rotor and the fuselage being discretized into a number of aerodynamic load calculation points in an embodiment of the present invention. In the above step 3, the unsteady stern flow field data is imported into each aerodynamic load calculation point to simulate the interference of the unsteady stern flow field on the helicopter. The specific implementation process of step 3 may include:

[0120] Step 31: Divide one cycle of the blade movement into N P The station is set at the station corresponding to time t0, and the instantaneous velocity field of the unsteady stern flow field is W gust (t0,u G0 ,v G0 ,w G0 );

[0121] Step 32: The coordinates of one of the micro-segments of the blade in the blade coordinate system are (t0, x b0 ,y b0 ,z b0 ), then the coordinates of the blade micro-segment in the body coordinate system are converted to:

[0122]

[0123] The coordinates of the blade micro-segment converted into the ship coordinate system are:

[0124]

[0125] Among them, T Rb represents the transformation matrix from the blade coordinate system to the rotor rotating coordinate system, T HR represents the transformation matrix from the rotor rotating coordinate system to the hub coordinate system, TBH represents the transformation matrix from the hub coordinate system to the body coordinate system, T EB represents the transformation matrix from the body coordinate system to the inertial coordinate system, T CE represents the transformation matrix from the inertial coordinate system to the ship coordinate system; (x MR ,y MR ,z MR ) is the coordinate of the rotor hub center;

[0126] Step 33, interpolate the unsteady stern flow field data obtained in step 1 using the distance derivative weighted method to obtain the disturbance velocity component (t0, u GC0 ,v GC0 ,w GC0 ), and transform the velocity component to the blade coordinate system. After the transformation, the instantaneous velocity component of the micro-segment center is:

[0127]

[0128] Among them, T bR , T RH , T HB , T BE and T EC They are the inverse matrices corresponding to the above coordinate transformation matrices respectively;

[0129] Step 34, repeating steps 32 to 33 to obtain instantaneous velocity components of the centers of other micro-segments on the blade;

[0130] Step 35, the blade phase angle is advanced forward by one station, and the time remains unchanged, and steps 32 to 34 are repeated until the blade completes one rotational motion, and the instantaneous velocity component of each micro-segment in the blade at each station is obtained; and each instantaneous velocity component obtained is calculated and averaged respectively, so as to obtain the average unsteady stern flow field velocity component in the propeller disk plane at this moment:

[0131]

[0132] in, It represents the instantaneous velocity component obtained by interpolation of the Ith micro-segment on the blade at the K station; the average unsteady stern flow field velocity component is used to calculate the control margin of the coaxial rigid rotor helicopter.

[0133] Furthermore, step 3 of the embodiment of the present invention may also include the following steps:

[0134] Step 36, according to the calculation results of steps 34 and 35, the aerodynamic force and moment of the blade at time t0 are obtained, and steps 32 to 34 are repeated to obtain the aerodynamic force and moment of each blade, and the aerodynamic force and moment of all blades are summed respectively to obtain the aerodynamic force and moment of the entire rotor at time t0;

[0135] Step 37, the blade phase angle is advanced forward by one station and the time is advanced forward by Δt, and the unsteady stern flow field data of the blade micro-segment at the time t0+Δt is interpolated to obtain the instantaneous velocity component of the center of the blade micro-segment at the time t0+Δt;

[0136] Step 38, by repeating steps 32 to 34, and steps 36 to 37, the instantaneous velocity components of each micro-segment in each blade at each position and at each moment are obtained, thereby simulating the unsteady interference of the coaxial rigid rotor with the turbulent vortex structure during the periodic motion.

[0137] The landing flight characteristics analysis method for a coaxial rigid rotor helicopter provided in an embodiment of the present invention has three main core contents: high-precision unsteady stern flow field data acquisition, establishment of coaxial rigid rotor helicopter flight dynamics modeling that can couple unsteady stern flow field data, and data transmission strategy from unsteady stern flow field data to the flight dynamics model. In the technical solution of the embodiment of the present invention, the high-precision unsteady stern flow field obtained in the first part can be imported into the coaxial rigid rotor helicopter flight dynamics model established in the second part; the "data transmission strategy based on unidirectional coupling" in the third part is the key innovation of the present invention. In this part of the technology, the present invention adopts the idea of ​​"nested grids" to discretize the blade rotation motion into N P At each station, the unsteady interference between the coaxial rigid rotor and the large-scale vortex of the stern flow is achieved by matching the blade motion with the unsteady stern flow field on a time scale.

[0138] The landing flight characteristics analysis method for a coaxial rigid rotor helicopter provided in the embodiment of the present invention can quickly and accurately provide the changing characteristics of the control margin and the pilot workload during the landing of the coaxial rigid rotor helicopter. In addition, the simulation cycle for a landing process in the embodiment of the present invention is only about 10 days. Compared with the sea test flight, the present invention can significantly shorten the test cycle, save manpower and material resources, and reduce risks.

[0139] The following is a schematic illustration of a specific embodiment of the method for analyzing landing flight characteristics of a coaxial rigid rotor helicopter provided in an embodiment of the present invention.

[0140] The core contents of the landing flight characteristics analysis method for a coaxial rigid rotor helicopter provided in this specific implementation example mainly include three parts: high-precision unsteady stern flow field data acquisition, establishment of coaxial rigid rotor helicopter flight dynamics modeling that can couple unsteady stern flow field data, and data transmission strategy of unsteady stern flow field data to flight dynamics model. The specific implementation methods of the above three parts are described in detail below.

[0141] (1) Obtaining high-confidence unsteady stern flow field data based on the detached vortex method

[0142] This specific embodiment uses a high-precision detached vortex method to perform numerical calculations on the unsteady stern flow field. The calculation method is: use grid generation software to generate a rectangular calculation domain grid of 10Ls (length) × 8Ls (width) × 6Ls (height); Ls is the length of the ship. The inlet boundary and outlet boundary of the rectangular calculation domain grid are set as velocity inlet conditions and pressure outlet conditions respectively; the hull is a no-slip wall, and the sea level and other calculation domain surfaces are set as slip walls. Since the boundary layer grid of the ship has little effect on the stern flow characteristics, under various control conditions, the thickness of the first layer of grids in the boundary layer grid (referring to the innermost layer of the boundary layer grid) is 1mm to 2mm (in this specific embodiment, the thickness of the first layer of grids is, for example, 1.5mm) to ensure that the y of the wall function used in the turbulence model calculation is met. + Condition(y + The value range is 28 to 300), and the boundary layer grid growth rate is 1.05 to 1.15 (in this specific embodiment, the growth rate is 1.1, for example). The grid size (Δ0 / H) near the ship is about 3.6×10 -2 , in order to ensure the fine capture of the turbulent flow near the ship, the total number of grids in the final ship calculation model is 7.5×10 6 .

[0143] It should be noted that y + is calculated by the first layer grid thickness and the local speed of the ship. Due to the different local speeds, y + The value is a range.

[0144] Through the above calculation and setting values ​​of this step, high-precision unsteady stern flow field data of a preset time (e.g., 5s to 10s) can be obtained to obtain a stable unsteady stern flow field topology structure; thereby, the unsteady stern flow field data of a specified time (e.g., 30s) can be calculated, and 7-dimensional unsteady stern flow field data can be output.

[0145] (2) Constructing the flight dynamics model of a coaxial rigid rotor helicopter

[0146] Unlike conventional single-rotor helicopters, coaxial rigid-rotor helicopters use a coaxially reversed layout of upper and lower rotors, which results in significant aerodynamic interference between the upper and lower rotors during flight, and the interference characteristics change continuously with the flight state.

[0147] 2.1, determine the upper and lower rotor interference factors.

[0148] In this specific embodiment, the concept of interference factor (δ1, δ2) is introduced to quantitatively calculate the mutual interference strength between the upper and lower rotors, and its value can be determined by the slipstream theory. As shown in FIG2, the specific situation of the mutual interference between the upper rotor and the lower rotor in a coaxial rigid rotor helicopter based on the slipstream theory is shown. In this specific embodiment, the rotor disc area is defined as S, such as Figure 2a As shown, the interference area between the upper rotor wake and the lower rotor blade is S′1( Figure 2a The shaded area in the figure), the interference area between the lower rotor induced flow and the upper rotor blade is S′2 ( Figure 2b The shaded area in the figure). The average induced velocities in the lower rotor and upper rotor disc planes are v1 and v2 respectively. Assume that the forward flight speed of the helicopter is V x , the distance between the lower and upper rotors is d, and the trail inclination angles of the lower and upper rotors are χ1 and χ2 respectively. According to the Biot-Savart theorem, the interference factors between the upper and lower rotors can be determined as:

[0149] δ1=S′1 / S f(d,χ2);

[0150] δ2=S′2 / S f(-d,χ1);

[0151] 2.2, determine the inflow model of the upper and lower rotors.

[0152] In this specific embodiment, the Pitt-Prters static non-uniform inflow model in the form of a first-order harmonic wave is used to describe the induced velocity distribution of the upper and lower rotors to ensure sufficient accuracy.

[0153] In this specific embodiment, by introducing interference factors, the inflow models of the upper rotor and the lower rotor are determined as follows:

[0154]

[0155]

[0156] Among them, r is the distance from the blade micro-segment to the center of the hub, R is the blade radius, K1 and K2 are the induced velocity distortion factors, ψ1 and ψ2 are the phase angles of the lower rotor and upper rotor blades respectively.

[0157] 2.3, determine the aerodynamic loads on the upper and lower rotors.

[0158] In this specific embodiment, in order to consider the influence of the unsteady stern flow field on the rotor aerodynamic load, the process of rotor aerodynamic modeling includes the following steps:

[0159] Step a, divide a single blade into several micro-segments, and import the unsteady stern flow field data at the center of each micro-segment when conducting aircraft / ship dynamic interface research. Considering that the outer side of the blade is the main source of lift, the length of the micro-segment gradually decreases from the blade root to the blade tip, and the area of ​​the ring determined by the inner and outer boundaries of each micro-segment in the propeller disc plane is guaranteed to be the same. The schematic diagram of the blade micro-segment division is shown in the figure below: Figure 3 shown.

[0160] like Figure 3 The dimensionless flapping hinge offset of the blade micro-segment shown is ε = e / R, ξ represents the dimensionless distance from the blade root cut to the flapping hinge, c T and c R are the chord lengths at the blade tip and root cut, respectively. Then the distance from the center of the first blade micro-segment starting from the blade root to the flapping hinge can be:

[0161]

[0162] Where N SS is the number of blade micro-segments.

[0163] The distances from the center of other blade micro-segments to the flapping hinge are:

[0164] Step b, taking the upper rotor as an example, the tangential velocity, spanwise velocity and normal velocity component U of the center of the upper rotor micro-segment T2 , U R2 and U P2 They are:

[0165]

[0166]

[0167] U R2 =[μ 02 cosψ2cos(β2+θ pre )+λ 02 sin(β2+θ pre )]ΩR+v Gb ;

[0168] In the formula, μ 02 , 02 are the upper rotor advance ratio and inflow ratio, p W2 ,q W2 are the pitch and roll angular velocities of the rotor under the wind axis, β2, are the blade flapping angle and flapping angular velocity, θpre is the rotor pre-cone angle, u Gb 、u Gb 、u Gb They respectively represent the relative motion velocity components of the blade micro-segment caused by the stern flow.

[0169] Step c, using the above velocity components to calculate the corresponding micro-segment aerodynamic force, and accumulating the aerodynamic forces of each micro-segment to obtain the aerodynamic force at the root of a single blade, as shown in the following formula:

[0170]

[0171] And the torque generated on the flapping hinge and the swinging vibration can be obtained as:

[0172]

[0173] In the formula, F PI2 、F TI2 、F RI2 are the normal force, tangential force and span force of the blade micro-segment I respectively, and y2 represents the distance from the center of the blade micro-segment to the equivalent flapping hinge.

[0174] Step d, summing up the aerodynamic forces and moments of each blade to obtain the aerodynamic load of the upper rotor or the lower rotor.

[0175] 2.4, determine the upper and lower rotor flapping motion models respectively.

[0176] In this specific embodiment, the coaxial rigid rotor blade has no flapping hinge, but still has flapping motion. The first-order flapping expression of the coaxial rigid rotor is:

[0177]

[0178] Where, e is the equivalent swing offset, M β ,I β are the blade mass static moment and inertia moment respectively, K0 is the blade flapping stiffness when not rotating, Ω is the rotor speed, ω n is the first-order flapping frequency of the coaxial rigid rotor.

[0179] In this specific embodiment, a method based on the equivalent flapping extension and equivalent flapping constraint spring stiffness of the first-order elastic flapping vibration mode is used to establish a flapping motion model. The basic idea is:

[0180] Step a, use the first-order rigid flapping motion of the equivalent articulated rotor to fit the first-order elastic flapping motion of the rigid rotor. The equivalent flapping hinge offset is the intersection of the outer end tangent of the first-order elastic flapping vibration mode and the x-axis (the positive direction of the x-axis is from the blade root to the blade tip), such as Figure 4 shown.

[0181] In addition, in order to ensure that the flapping frequency of the coaxial rigid rotor is the same as that of the equivalent articulated rotor, the stiffness of the constrained torsion spring applied at the equivalent flapping hinge is:

[0182]

[0183] Through the above method, a flapping motion model of the coaxial rigid rotor can be established.

[0184] Step b: During the blade movement, the moments acting on the equivalent flapping hinge are mainly: centrifugal moment M CF , swinging inertia moment M I , aerodynamic moment M A 、Coriolis moment M caused by the angular velocity of the helicopter Cor 、Inertia moment M caused by fuselage angular acceleration BA , M caused by the body acceleration BL , blade gravity moment M R , and the torque M generated by the torsion spring R , whose expressions are:

[0185] M CF =Ω 2 (I β cosβ2+eM β )sinβ2;

[0186]

[0187] M Cor =-2Ω(I β +eM β )(p W2 cosψ2-q W2 sinψ2);

[0188]

[0189]

[0190] M G =M β g;

[0191]

[0192] M R =K β β2;

[0193] in, is the blade flapping angular acceleration, are the pitch and roll angular accelerations of the rotor in the wind axis system, (u, v, w) are the velocity components of the center of gravity of the body in the body coordinate system, are the acceleration of the body's center of gravity in the body coordinate system. The flapping motion model of the coaxial rigid rotor established in this step is:

[0194]

[0195] Among them, C1, C2, and C3 are the damping matrix, stiffness matrix, and excitation force matrix of the rotor system, respectively.

[0196] 2.5, determine the aerodynamic model of the fuselage, horizontal tail, vertical tail and tail rotor respectively, the aerodynamic model includes aerodynamic force and moment;

[0197] 2.6, determine the flight dynamics model of the coaxial rigid rotor helicopter. The specific steps include:

[0198] Step a, synthesize the aerodynamic forces and moments of the above components at the center of gravity of the aircraft, and obtain the resultant force and moment of the helicopter center of gravity in the aircraft coordinate system (right-hand system), which are:

[0199]

[0200]

[0201] In the formula, on the left side of the equation: L is the rolling moment, M is the yaw moment, and N is the pitch moment; the subscripts X, Y, and Z represent the vector directions, the subscripts 1 and 2 represent the lower rotor and the upper rotor respectively, the subscripts MR, F, H, and V represent the rotor, fuselage, horizontal tail, and vertical tail respectively, Θ and Φ are the fuselage pitch angle and roll angle respectively, m is the total mass of the helicopter, and g is the acceleration of gravity.

[0202] In addition, Newton's second law and the momentum theorem can be used to obtain the rigid body dynamics models of the helicopter's center of gravity movement and its rotation around the center of gravity:

[0203]

[0204]

[0205] In the formula, I and ω * Represent the cross product matrix of the inertia matrix and angular velocity, (p, q, r) and are the angular velocity and angular acceleration of the body's center of gravity in the body coordinate system respectively.

[0206] Step b, combining the rotor flapping motion model with the helicopter rigid body dynamics model, can obtain the flight dynamics model of the coaxial rigid rotor helicopter, and its simplified form is:

[0207]

[0208] Among them, t is the time variable, X and U are the helicopter state variable and manipulated variable respectively, and W gust is the relative speed of helicopter components caused by the unsteady stern flow field.

[0209] (3) a data transmission strategy based on one-way coupling, i.e., a data transmission strategy of the unsteady stern flow field data to the flight dynamics model; specifically, coupling the unsteady stern flow field data to the flight dynamics model of the coaxial rigid rotor helicopter to calculate the control margin and pilot workload of the coaxial rigid rotor helicopter;

[0210] In this specific embodiment, a data transmission strategy is established based on the idea of ​​"one-way coupling", that is, only the interference of the unsteady stern flow on the helicopter is considered. In this method, the rotor and the fuselage are discretized into several calculation points of aerodynamic loads, such as Figure 5 shown.

[0211] In this specific embodiment, an example is given in which there are 10 aerodynamic load calculation points on each blade, one each on the fuselage, horizontal tail, and vertical tail, for a total of 83 aerodynamic load calculation points. The interference of the unsteady stern flow field on the helicopter is simulated by importing the unsteady stern flow field data into each aerodynamic load calculation point.

[0212] Before data transmission, an effective unsteady stern flow field must be obtained, and its range should include all areas that the aerodynamic center of each component of the helicopter passes through during the landing flight. Figure 6 FIG. 1 is a schematic diagram of the stern flow field data domain established for the lateral displacement process in an embodiment of the present invention. The unsteady stern flow field data W obtained by CFD calculation gust (t,u,v,w) cannot be directly used in the helicopter flight mechanics model. Its storage method must be reorganized in the form of a structural grid to facilitate the interpolation of the aerodynamic centers of the helicopter components. When analyzing the balance characteristics of helicopter landing, it is also necessary to perform time-average processing on the unsteady stern flow field data to obtain the time-averaged flow field data with the same storage structure.

[0213] The rotor blades are waving periodically while rotating around the hub, and their spatial positions in the unsteady stern flow field will also change due to the change in the body attitude, which makes the transmission process of CFD data to the rotor aerodynamic model more complicated. Here, the interpolation process of the blade micro-segment is used to explain this in detail. This step includes the following specific implementation process:

[0214] 3.1, to simulate the periodic motion of the blade, one cycle of the blade motion is divided into N P Assume that at time t0, the instantaneous velocity field of the unsteady stern flow field is W gust (t0,u G0 ,v G0,w G0 );

[0215] 3.2, the coordinates of one of the micro-segments of the blade in the blade coordinate system are (t0,x b0 ,y b0 ,z b0 ), the coordinates of the blade micro-segment in the body coordinate system can be converted to:

[0216]

[0217] The coordinates of the blade micro-segment in the ship coordinate system are further converted to:

[0218]

[0219] Among them, T Rb represents the transformation matrix from the blade coordinate system to the rotor rotating coordinate system, T HR represents the transformation matrix from the rotor rotating coordinate system to the hub coordinate system, T BH represents the transformation matrix from the hub coordinate system to the body coordinate system, T EB represents the transformation matrix from the body coordinate system to the inertial coordinate system, T CE represents the transformation matrix from the inertial coordinate system to the ship coordinate system; (x MR ,y MR ,z MR ) are the coordinates of the rotor hub center.

[0220] 3.3, after obtaining the position coordinates of the blade micro-segment in the ship coordinate system, the distance derivative weighted method is used to interpolate the unsteady stern flow field data, as shown in the following formula:

[0221]

[0222] in,

[0223] Where s represents the velocity component to be interpolated at the center of the blade micro-segment, s i is the velocity component of the data point adjacent to the micro-segment center, (x Ci ,y Ci ,z Ci ) is the position coordinate of the adjacent point, d i Indicates the distance from the adjacent point to the micro-segment center, α is the exponent for adjusting the inverse of the distance weight, and n is the number of adjacent data points. In this study, the unsteady stern flow field data is stored in the form of a structured grid, and the blade micro-segment center will eventually fall into a rectangular unit in the data domain, so n = 8. Practice has shown that α = 2 can guarantee a high interpolation accuracy. The above interpolation method can be used to obtain the instantaneous disturbance velocity component (t0, u GC0,v GC0 ,w GC0 ); Then, the velocity component is converted to the blade coordinate system. After the conversion, the instantaneous velocity component of the micro-segment center is:

[0224]

[0225] T bR , T RH , T HB , T BE and T EC are the inverse matrices of the above coordinate transformation matrices respectively.

[0226] 3.4, performing steps 3.2 to 3.3 on other micro-segments on the blade to obtain instantaneous velocity components of the centers of other micro-segments on the blade;

[0227] 3.5, (this step is performed only once), move the blade phase angle forward by one station (time remains unchanged), repeat steps 3.2 to 3.4 until the blade completes a rotation. Sum the obtained velocity components and then calculate their average values ​​to obtain the average stern flow field velocity component (t0, u GAV0 ,v GAV0 ,w GAV0 ), as shown below:

[0228]

[0229] in, represents the instantaneous velocity component obtained by interpolation of the Ith micro-segment on the blade at the K station; and the average unsteady stern flow field velocity component calculated by this formula is used to calculate the control margin of the coaxial rigid rotor helicopter.

[0230] 3.6. Based on the calculation results of steps 3.4 and 3.5, the aerodynamic force and moment of the blade at time t0 can be obtained; steps 3.2 to 3.4 are repeated for other blades to obtain the aerodynamic force and moment of each blade, and the aerodynamic force and moment of all blades are summed up to obtain the aerodynamic force and moment of the entire rotor at time t0.

[0231] 3.7, the blade moves forward one station. On the one hand, the spatial position of the blade micro-segment is updated, and on the other hand, the time is simultaneously advanced by Δt. At this time, the stern flow field data of the blade micro-segment at time t0+Δt is interpolated, and steps 3.2~3.4, 3.6~3.7 are repeated to simulate the unsteady interference between the rotor blade and the turbulent vortex structure during the periodic motion.

[0232] It should be noted that the interpolation process of the aerodynamic centers of other parts of the helicopter is similar to that of the rotor, but these aerodynamic centers are all on the fuselage, and only involve the coordinate transformation between the fuselage coordinate system, the inertial coordinate system and the ship coordinate system. The data transmission process is relatively simple and will not be repeated here. Through the research content of the embodiment of the present invention, the high-precision ship unsteady stern flow field obtained in the first part can be introduced into the coaxial rigid rotor helicopter flight dynamics model established in the second part.

[0233] In addition, the "data transmission strategy based on unidirectional coupling" in the third part is the key innovation of the present invention. In this part of the technology, the present invention adopts the idea of ​​"nested grid" to discretize the blade rotation motion into N P At each station, the unsteady interference between the coaxial rigid rotor and the large-scale vortex of the stern flow is achieved by matching the blade motion with the unsteady stern flow field on a time scale.

[0234] The landing flight characteristics analysis method for a coaxial rigid rotor helicopter provided by the embodiment of the present invention can quickly and accurately provide the changing characteristics of the control margin and the pilot workload during the landing process of the coaxial rigid rotor helicopter. Figure 7 As shown in FIG. 1 , it is a schematic diagram of the curve of the change of the control amount, attitude angle and required power when the coaxial rigid rotor helicopter performs lateral displacement at different heights in the embodiment of the present invention. It can be seen that within the range of -0.5≤y / B≤0.0, the helicopter's total distance gradually increases. Figure 7 This is because when the helicopter moves toward the center of the deck, the upper and lower rotor thrusts are reduced due to the low-speed recirculation area and strong downwash in the landing area, and the pilot must increase the collective pitch input to maintain the stability of the helicopter altitude. In addition, the longitudinal cyclic pitch variation increases significantly within this range, as shown in Figure 7 As shown in Figure b, this is because the front half of the rotor is relatively strongly affected by the downwash, which increases the asymmetry of the front and rear thrust of the rotor, thereby forming a nose-down moment on the helicopter. The pilot also needs to increase the longitudinal cyclic pitch input to make the rotor disc fall back to offset the impact of this nose-down moment.

[0235] like Figure 8 As shown, it is a schematic diagram of the variation curve of the unsteady load level of the coaxial rigid rotor helicopter at different lateral displacement path heights in an embodiment of the present invention. On the whole, as the coaxial helicopter moves toward the center of the deck, the pilot workload gradually increases. This is because the free flow is blocked by the ship's superstructure and forms a low-speed recirculation zone behind the hangar. At the same time, the turbulent vortex structure generated by its flow separation will also form a strong turbulent area in the landing area. During the lateral displacement process, the helicopter gradually enters the strong unsteady flow field environment from the uniform flow area, which increases the unsteady load level of the helicopter.

[0236] The above calculation results are consistent with the actual flight experience of pilots, indicating that the method established by the present invention has good accuracy and good practicability.

[0237] In addition, the simulation period for a landing process in the embodiment of the present invention is only about 10 days. Compared with sea trials and flights, the present invention can significantly shorten the test period, save manpower and material resources, and reduce risks.

[0238] Although the embodiments disclosed in the present invention are as above, the contents described are only embodiments adopted to facilitate understanding of the present invention and are not intended to limit the present invention. Any technician in the field to which the present invention belongs can make any modifications and changes in the form and details of implementation without departing from the spirit and scope disclosed in the present invention, but the patent protection scope of the present invention shall still be subject to the scope defined in the attached claims.

Claims

1. A landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter, characterized in that: include: Step 1, using a high-precision detached vortex method to obtain high-confidence unsteady stern flow field data; Step 2, based on the unsteady stern flow field data, construct a flight dynamics model of a coaxial rigid rotor helicopter for coupling the unsteady stern flow field data; Step 3, coupling the unsteady stern flow field data to the flight dynamics model of the coaxial rigid rotor helicopter to calculate the control margin and pilot workload of the coaxial rigid rotor helicopter; Wherein, the coaxial rigid rotor helicopter adopts a layout form in which the upper rotor and the lower rotor are coaxial and reversed; the step 2 comprises: Step 21, respectively determining the interference factors of the upper rotor and the lower rotor, so as to obtain and quantitatively calculate the mutual interference intensity between the upper rotor and the lower rotor; Step 22, determining the inflow models of the upper rotor and the lower rotor respectively according to the interference factor, so as to represent the induced velocity distribution of the upper rotor and the lower rotor; Step 23, determining the aerodynamic loads of the upper rotor and the lower rotor respectively; Step 24, respectively determining the flapping motion models of the upper rotor and the lower rotor; Step 25, respectively determining aerodynamic models of the fuselage, horizontal tail, vertical tail and tail rotor, wherein the aerodynamic models include aerodynamic forces and moments; Step 26, determining the flight dynamics model of the coaxial rigid rotor helicopter; The step 23 comprises: A single blade is divided into multiple micro-segments, and the unsteady stern flow field data is imported into the center of each micro-segment to obtain the tangential velocity component, spanwise velocity component and normal velocity component of each micro-segment center, and each velocity component is used to calculate the corresponding micro-segment aerodynamic force, and the aerodynamic force of each micro-segment is accumulated to obtain the aerodynamic force and moment of the root of the single blade; the aerodynamic force and moment of each blade in the upper rotor or the lower rotor are summed respectively to obtain the aerodynamic load of the upper rotor or the lower rotor; The step 24 comprises: a. The first-order rigid flapping motion of the equivalent articulated rotor is used to fit the first-order elastic flapping motion of the coaxial rigid rotor; the equivalent flapping hinge offset is the intersection of the outer end tangent of the first-order elastic flapping vibration mode and the x-axis, and the constraint torsion spring stiffness applied at the equivalent flapping hinge is required to be: Wherein, the x-axis is defined as the blade span direction, e is the equivalent flapping offset, M β ,I β are the blade mass static moment and inertia moment respectively, Ω is the rotor speed, ω n is the first-order flapping frequency of the coaxial rigid rotor; b, according to the centrifugal moment M CF , swinging inertia moment M I , aerodynamic moment M A 、Coriolis moment M caused by the angular velocity of the helicopter Cor 、Inertia moment M caused by fuselage angular acceleration BA , M caused by the body acceleration BL , blade gravity moment M R , and the torque M generated by the torsion spring R , the algebraic sum at the equivalent flapping hinge should be zero, and a flapping motion model of the coaxial rigid rotor is established.

2. The landing flight characteristics analysis method for a coaxial rigid rotor helicopter according to claim 1, characterized in that: The step 1 comprises: Step 11, generating a rectangular computational domain grid with a preset length, width, and height according to the shape and size of the ship; Step 12, setting the inlet boundary of the rectangular computational domain grid as the velocity inlet condition, and setting the outlet boundary as the pressure outlet condition; Step 13, setting the hull as a non-slip wall, and setting the sea level and other computational domain surfaces as sliding walls; Step 14, setting the thickness of the first layer of the boundary layer grid on the surface of the ship to 1 mm to 2 mm, and setting the boundary layer grid growth rate to 1.05 to 1.15, so as to obtain the total number of grids of the ship calculation model; Step 15, according to the calculation and setting values ​​of steps 11 to 14, a high-precision detached vortex method is used to calculate the unsteady stern flow field data of a preset time to obtain a stable unsteady stern flow field topology structure; Step 16, after obtaining the stable unsteady stern flow field topology structure, calculate the unsteady stern flow field data at the specified time, and output the 7-dimensional unsteady stern flow field data.

3. The landing flight characteristics analysis method for a coaxial rigid rotor helicopter according to claim 1, characterized in that: The step 21 comprises: The interference factors of the upper rotor and the lower rotor are deduced by the slipstream theory: δ1=S1′ / Sf(d,χ2); δ2=S2′ / Sf(-d,χ1); Where S is the rotor disc area, S1′ is the interference area between the upper rotor wake and the lower rotor disc, S2′ is the interference area between the lower rotor induced flow and the upper rotor disc, V x is the forward flight speed of the helicopter, d is the distance between the upper rotor and the lower rotor, χ1 and χ2 are the wake inclination angles of the lower rotor and the upper rotor respectively.

4. The landing flight characteristics analysis method for a coaxial rigid rotor helicopter according to claim 3, characterized in that: The step 21 comprises: The step 22 comprises: According to the interference factor in step 21, the inflow models of the upper rotor and the lower rotor are determined as follows: Among them, r is the distance from the blade micro-segment to the center of the hub, R is the blade radius, K1 and K2 are the induced velocity distortion factors, ψ1 and ψ2 are the phase angles of the lower rotor and upper rotor blades respectively.

5. The landing flight characteristics analysis method applicable to a coaxial rigid rotor helicopter according to any one of claims 1 to 4, characterized in that: The step 26 comprises: a. The aerodynamic forces and moments of all components are synthesized at the center of gravity of the fuselage to obtain the resultant force and moment acting on the center of gravity of the helicopter in the fuselage coordinate system, and the rigid body dynamics model of the movement of the center of gravity of the helicopter and its rotation around the center of gravity is obtained; b. Based on the flapping motion model and rigid body dynamics model of the coaxial rigid rotor, the flight dynamics model of the coaxial rigid rotor helicopter is obtained.

6. The landing flight characteristics analysis method for a coaxial rigid rotor helicopter according to claim 5, characterized in that: Each micro-segment of each blade of the upper rotor and the lower rotor, as well as the fuselage, the horizontal tail, the vertical tail and the tail rotor are respectively an aerodynamic load calculation point. In step 3, the unsteady stern flow field data is imported into each aerodynamic load calculation point to simulate the interference of the unsteady stern flow field on the helicopter; The step 3 specifically includes: Step 31: Divide one cycle of the blade movement into N P The station is set at the station corresponding to time t0, and the instantaneous velocity field of the unsteady stern flow field is W gust (t0,u G0 ,v G0 ,w G0 ); Step 32: The coordinates of one of the micro-segments of the blade in the blade coordinate system are (t0, x b0 ,y b0 ,z b0 ), then the coordinates of the blade micro-segment in the body coordinate system are: The coordinates of the blade micro-segment converted into the ship coordinate system are: Among them, T R b represents the transformation matrix from the blade coordinate system to the rotor rotating coordinate system, T HR represents the transformation matrix from the rotor rotating coordinate system to the hub coordinate system, T BH represents the transformation matrix from the hub coordinate system to the body coordinate system, T EB represents the transformation matrix from the body coordinate system to the inertial coordinate system, T CE represents the transformation matrix from the inertial coordinate system to the ship coordinate system; (x MR ,y MR ,z MR ) is the coordinate of the rotor hub center; Step 33, interpolate the unsteady stern flow field data obtained in step 1 using the distance derivative weighted method to obtain the disturbance velocity component (t0, u GC0 ,v GC0 ,w GC0 ), and transform the velocity component to the blade coordinate system. After the transformation, the instantaneous velocity component of the micro-segment center is: Among them, T bR , T RH , T HB , T BE and T EC They are the inverse matrices corresponding to the above coordinate transformation matrices respectively; Step 34, repeating steps 32 to 33 to obtain instantaneous velocity components of the centers of other micro-segments on the blade; Step 35, the blade phase angle is advanced forward by one station, and the time remains unchanged, and steps 32 to 34 are repeated until the blade completes one rotational motion, and the instantaneous velocity component of each micro-segment in the blade at each station is obtained; and each instantaneous velocity component obtained is calculated and averaged respectively, so as to obtain the average unsteady stern flow field velocity component in the propeller disk plane at this moment: Among them, N P N is the number of stations divided by one circle of blade movement. SS is the number of blade micro-segments, It represents the instantaneous velocity component obtained by interpolation of the Ith micro-segment on the blade at the K station; the average unsteady stern flow field velocity component is used to calculate the control margin of the coaxial rigid rotor helicopter.

7. The landing flight characteristics analysis method for a coaxial rigid rotor helicopter according to claim 6, characterized in that: The step 3 also includes: Step 36, according to the calculation results of steps 34 and 35, the aerodynamic force and moment of the blade at time t0 are obtained, and steps 32 to 34 are repeated to obtain the aerodynamic force and moment of each blade, and the aerodynamic force and moment of all blades are summed respectively to obtain the aerodynamic force and moment of the entire rotor at time t0; Step 37, the blade phase angle is advanced forward by one station and the time is advanced forward by Δt, and the unsteady stern flow field data of the blade micro-segment at the time t0+Δt is interpolated to obtain the instantaneous velocity component of the center of the blade micro-segment at the time t0+Δt; Step 38, by repeating steps 32 to 34, and steps 36 to 37, the instantaneous velocity components of each micro-segment in each blade at each position and at each moment are obtained, thereby simulating the unsteady interference of the coaxial rigid rotor with the turbulent vortex structure during the periodic motion.

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