Real-time resolving method and device for propeller spacing of coaxial double rotors

Abstract

The invention provides a real-time resolving method and device for the pitch of coaxial double rotors, and the method comprises the steps: 1, building a mathematic relation among the bending moment load-structural strain-displacement response of a blade section through the analysis and calculation information in a design stage; 2, selecting a proper blade section, and solving the displacement response at the position R of the blade tip according to the flight actual measurement load data and the mathematical relationship between the blade section bending moment load-structural strain-displacement response; 3, respectively calculating displacement responses of the upper rotor and the lower rotor according to blade profile flight actual measurement load data M (r, t) of the upper rotor and the lower rotor; and according to the displacement response of the upper and lower rotors and the static propeller spacing of the upper and lower rotors, calculating the propeller spacing of the upper and lower rotors.

Description

Technical Field

[0001] The invention belongs to the technical field of helicopter rotors, and in particular relates to a method and a device for real-time calculation of blade pitch of a coaxial twin rotor. Background Art

[0002] In order to reduce rotor drag, the installation distance between the upper and lower rotors of coaxial twin-rotor helicopters is reduced. However, the smaller installation distance between the rotors may cause blade collisions during complex flight, endangering flight safety. Therefore, it is necessary to measure and monitor the distance and change trend between the upper and lower blades in real time during flight and effectively control them to prevent blade collisions.

[0003] However, the problem of blade tip displacement response and blade pitch involves multiple disciplines such as flight mechanics and rotor dynamics. It is not feasible to directly simulate and analyze all flight conditions accurately. Therefore, reliable measurement and monitoring must be carried out in engineering practice to ensure flight safety. Due to the influence of the strong centrifugal force field of the rotor, limited installation space and high vibration environment of the whole aircraft, there are significant defects in the engineering feasibility, measurement accuracy and test data credibility of adding radar, optical measurement and other equipment to the rotor blades or fuselage structure to measure blade displacement response and blade pitch. Summary of the invention

[0004] The purpose of the present invention is to propose a method for calculating the tip pitch in real time based on measured data, and to solve the blade deformation and pitch information of the upper and lower rotors in all azimuth angles in the entire circumference through the mathematical relationship between section bending moment, structural strain and deformation.

[0005] In a first aspect, the present application provides a method for real-time calculation of the pitch of a coaxial twin-rotor, the method comprising:

[0006] Step 1: Using the analytical calculation information from the design phase, establish the mathematical relationship between the blade section bending moment load-structural strain-displacement response;

[0007] Step 2: Select a suitable blade profile and solve the displacement response at the blade tip R based on the flight measured load data and the mathematical relationship between the blade profile bending moment load-structural strain-displacement response;

[0008] Step 3: According to the flight measured load data M(r, t) of the blade profiles of the upper and lower rotors, the displacement responses of the upper and lower rotors are calculated respectively; according to the displacement responses of the upper and lower rotors and the static blade pitches of the upper and lower rotors, the blade pitches of the upper and lower rotors are calculated.

[0009] Furthermore, step 1 includes:

[0010] Step 11: Establish the flapping direction displacement response Z(x, t) of the blade at time t and spanwise position x:

[0011]

[0012] In the above formula, γ n is the vibration mode of each order of flapping mode, q n is the corresponding generalized modal displacement;

[0013] Step 12: According to the flapping direction displacement response Z(x, t), the blade section bending moment load M(x, t) at time t and spanwise position x is obtained:

[0014]

[0015] Among them, EI flap (x) is the flapping stiffness, γ″ n (x) is the structural strain.

[0016] Furthermore, step 2 includes:

[0017] Step 21: Select the flight measured load data M(r,t) of the appropriate blade profile r:

[0018] M(r,t)=Mb1(r)·q1(t)+Mb2(r)·q2(t)+Mb3(r)·q3(t)+…;

[0019] Step 22: Calculate the displacement response Z(R, t) at the blade tip R:

[0020] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t)+…;

[0021] Step 23: Obtain the first-order generalized modal displacement q1(t) through the flight measured load data M(r,t):

[0022]

[0023] Step 24: Substitute q1(t) into the displacement response Z(R, t) to obtain the displacement response at the blade tip R;

[0024]

[0025] Furthermore, step 2 includes:

[0026] Step 21: Select the flight measured load data of three blade sections with spanwise position coordinates a, b, and c: M(a, t), M(b, t), and M(c, t):

[0027] M(a,t)=Mb1(a)·q1(t)+Mb2(a)·q2(t)+Mb3(a)·q3(t);

[0028] M(b,t)=Mb1(b)·q1(t)+Mb2(b)·q2(t)+Mb3(b)·q3(t);

[0029] M(c,t)=Mb1(c)·q1(t)+Mb2(c)·q2(t)+Mb3(c)·q3(t);

[0030] Step 22: Solve the equations of M(a, t), M(b, t) and M(c, t) to obtain q1(t), q2(t) and q3(t);

[0031] Step 23: Perform a third-order truncation on the displacement response at the blade tip R, and substitute q1(t), q2(t), and q3(t) to obtain the blade tip displacement Z(R, t):

[0032] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t).

[0033] In a second aspect, the present application provides a real-time calculation device for the pitch of a coaxial twin-rotor, the device comprising a mathematical relationship establishment unit, a displacement response calculation unit and a pitch calculation unit, wherein:

[0034] A mathematical relationship building unit is used to establish the mathematical relationship between the blade section bending moment load-structural strain-displacement response using the analytical calculation information from the design phase;

[0035] The displacement response calculation unit is used to select a suitable blade profile and solve the displacement response at the blade tip R according to the flight measured load data and the mathematical relationship between the blade profile bending moment load-structural strain-displacement response;

[0036] The blade pitch calculation unit is used to calculate the displacement responses of the upper and lower rotors according to the flight measured load data M(r, t) of the blade profiles of the upper and lower rotors; and calculate the blade pitch of the upper and lower rotors according to the displacement responses of the upper and lower rotors and the static blade pitch of the upper and lower rotors.

[0037] Furthermore, the mathematical relationship establishing unit is specifically used for:

[0038] Establish the flapping direction displacement response Z(x, t) of the blade at time t and spanwise position x:

[0039]

[0040] In the above formula, γ n is the vibration mode of each order of flapping mode, q n is the corresponding generalized modal displacement;

[0041] According to the flapping direction displacement response Z(x, t), the blade section bending moment load M(x, t) at time t and spanwise position x is obtained:

[0042]

[0043] Among them, El flap (x) is the flapping stiffness, γ″ n (x) is the structural strain.

[0044] Furthermore, the displacement response calculation unit is specifically used for:

[0045] Select the flight measured load data M(r,t) of the appropriate blade profile r:

[0046] M(r,t)=Mb1(r)·q1(t)+Mb2(r)·q2(t)+Mb3(r)·q3(t)+…;

[0047] Calculate the displacement response Z(R, t) at the blade tip R:

[0048] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t)+…;

[0049] The first-order generalized modal displacement q1(t) is obtained through the flight measured load data M(r,t):

[0050]

[0051] Substitute q1(t) into the displacement response Z(R, t) to obtain the displacement response at the blade tip R;

[0052]

[0053] Furthermore, the displacement response calculation unit is specifically used for:

[0054] Select the flight measured load data of three blade sections with spanwise position coordinates a, b, and c: M(a, t), M(b, t), and M(c, t):

[0055] M(a,t)=Mb1(a)·q1(t)+Mb2(a)·q2(t)+Mb3(a)·q3(t);

[0056] M(b,t)=Mb1(b)·q1(t)+Mb2(b)·q2(t)+Mb3(b)·q3(t);

[0057] M(c,t)=Mb1(c)·q1(t)+Mb2(c)·q2(t)+Mb3(c)·q3(t);

[0058] Solve the system of equations M(a, t), M(b, t) and M(c, t) to obtain q1(t), q2(t) and q3(t);

[0059] The displacement response at the blade tip R is truncated to the third order, and q1(t), q2(t) and q3(t) are substituted to obtain the blade tip displacement Z(R, t):

[0060] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t).

[0061] To summarize, the present application provides a real-time calculation method for the propeller pitch of a coaxial twin-rotor, which monitors and measures the propeller pitch in real time during flight through the existing measured blade load information on the aircraft, realizes early warning of insufficient propeller pitch and provides reference information for flight control / manipulation to ensure flight safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0062] Figure 1 A schematic diagram of the forces acting on the blade within the flapping plane provided for this application;

[0063] Figure 2 Time domain information of blade flapping moment loads for two typical sections provided for this application;

[0064] Figure 3 A diagram of the time domain information of the blade tip displacement calculated based on the measured load of the blade provided in this application;

[0065] Figure 4 A diagram showing the spanwise distribution of the modal bending moment of a certain vibration mode provided in this application;

[0066] Figure 5 A diagram of the spanwise distribution of the second-order derivative of a certain vibration mode provided in this application;

[0067] Figure 6 This is a diagram of the time domain information of the propeller pitch obtained by solving the upper and lower blade displacements provided in this application. DETAILED DESCRIPTION

[0068] The flight measured blade profile load is the measurement information that is already available on the aircraft and has a relatively mature application. It includes the influence of steady-state factors such as the control balance in the rotor dynamics analysis, and the influence of the Coriolis force caused by transient factors in the professional research of flight dynamics. It is a true reflection of the rotor load. Based on the existing measured profile load information on the aircraft, the present invention solves the blade deformation and blade spacing information of the upper and lower rotors in all azimuths of the entire circumference through the mathematical relationship between profile bending moment-structural strain-deformation.

[0069] The rotor tip pitch of a coaxial twin-rotor helicopter is an important parameter that endangers flight safety. This method numerically solves the rotor tip pitch based on the measured blade load information, realizes real-time monitoring, and provides a reference for the safe flight of the helicopter.

[0070] In the modal space, the bending moment of the blade section is equal to the product of the modal bending moment and the generalized modal displacement, and the displacement response of the blade is equal to the product of the modal vibration shape and the generalized modal displacement. Both the modal bending moment and the modal vibration shape are known quantities. Therefore, the generalized modal displacement of the blade can be calculated by measuring the bending moment of the blade section, and then the displacement response of the blade and the blade tip can be calculated in combination with the modal vibration shape information.

[0071] There are three main types of inputs for the "modal displacement method": blade flapping stiffness information, flapping modal parameter information, and section bending moment information. In the design stage, the flapping modal parameter information and section bending moment information are both given by the structural design and dynamic design professionals. In the flight test stage, the blade flapping stiffness information and flapping modal parameter information are both known basic parameters for structural design and dynamic design, which are corrected by combining the stiffness test and dynamic test experimental data. The measurement of section bending moment information is a relatively mature testing technology. It can be directly measured in flight through patches, and then the real-time blade tip spacing can be solved by the "modal displacement method" calculation method to achieve real-time measurement and monitoring of blade tip spacing.

[0072] The present application provides a real-time calculation method for the blade pitch of a coaxial twin rotor based on the deformation and blade pitch of the upper and lower rotor blades of the flight measured load information, and the method includes:

[0073] Step 1: Using the analytical calculation information from the design phase, establish the mathematical relationship between the blade section bending moment load-structural strain-displacement response;

[0074] Specifically, step 1 includes:

[0075] Step 11: Establish the flapping direction displacement response Z(x, t) of the blade at time t and spanwise position x:

[0076]

[0077] In the above formula, γ n is the vibration mode of each order of flapping mode, q n is the corresponding generalized modal displacement;

[0078] Step 12: According to the flapping direction displacement response Z(x, t), the blade section bending moment load M(x, t) at time t and spanwise position x is obtained:

[0079]

[0080] Among them, EI flap (x) is the flapping stiffness, γ″n (x) is the structural strain.

[0081] like Figure 1 As shown, the modal bending moment Mb n (x) is the product of the flapping stiffness and the second-order derivative of the flapping vibration mode. The information of the flapping stiffness, flapping vibration mode and its derivative (strain term) is obtained in the analysis and calculation in the design stage. At this point, the mathematical relationship between the blade section bending moment load-structural strain-displacement response is established.

[0082] Step 2: Select a suitable blade profile and solve the displacement response at the blade tip R based on the flight measured load data and the mathematical relationship between the blade profile bending moment load-structural strain-displacement response;

[0083] According to the modal bending moment and strain distribution information, a set of appropriate blade profiles are selected, and the actual flight load data (such as Figure 2 Schematic diagram) Solve the algebraic equations between bending moment load-structural strain-deformation to obtain the deformation information of the blade (such as Figure 3 (indicated).

[0084] Under the action of a certain order of unit modal displacement, the maximum modal moment coefficient on a certain spanwise section means that the contribution of the modal component of this order to the swinging moment at this location is the largest, which is conducive to solving the displacement of this order mode (such as Figure 4 At the same time, the strain at the peak of the second-order derivative of this vibration mode is the largest, which is most conducive to the measurement of the strain / bending moment load at this location, that is, the accuracy of the swinging moment load measurement at this location is the highest (such as Figure 5 Schematic diagram). Therefore, when selecting the spanwise direction of the blade section, the contribution of the modal component and the measurement accuracy of the load must be comprehensively considered. Therefore, the section for moment measurement should be selected as much as possible, where the modal moment coefficient and the second-order derivative of the vibration mode are both near the peak point in the spanwise direction.

[0085] Specifically, if only the influence of the first-order flapping mode is considered, step 2 includes:

[0086] Step 21: Select the flight measured load data M(r,t) of the appropriate blade profile r:

[0087] M(r,t)=Mb1(r)·q1(t)+Mb2(r)·q2(t)+Mb3(r)·q3(t)+…(A)

[0088] Step 22: Calculate the displacement response Z(R, t) at the blade tip R:

[0089] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t)+…(B)

[0090] In the above two formulas, Mbn (r) and γ1(R) are known in the design stage. Simplifying formula (A),

[0091]

[0092] Step 23: Obtain the first-order generalized modal displacement q1(t) through the flight measured load data M(r,t):

[0093]

[0094] Step 24: Substitute q1(t) into the displacement response Z(R, t) to obtain the displacement response at the blade tip R.

[0095]

[0096] Specifically, if the influence of the first three flapping modes is considered, step 2 includes:

[0097] Step 21: Select the flight measured load data of three blade sections with spanwise position coordinates of a, b, and c, and write them as: M(a, t), M(b, t), and M(c, t), respectively;

[0098] According to formula (A), we can get:

[0099] M(a,t)=Mb1(a)·q1(t)+Mb2(a)·q2(t)+Mb3(a)·q3(t)

[0100] M(b,t)=Mb1(b)·q1(t)+Mb2(b)·q2(t)+Mb3(b)·q3(t)

[0101] M(c,t)=Mb1(c)·q1(t)+Mb2(c)·q2(t)+Mb3(c)·q3(t)

[0102] Step 22: Solve the equations of M(a, t), M(b, t) and M(c, t) to obtain q1(t), q2(t) and q3(t);

[0103] Step 23: Perform a third-order truncation on the displacement response at the blade tip R, and substitute q1(t), q2(t), and q3(t) to obtain the blade tip displacement Z(R, t):

[0104] Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t).

[0105] Step 3: According to the flight measured load data M(r, t) of the blade profiles of the upper and lower rotors, the displacement responses of the upper and lower rotors are calculated respectively; according to the displacement responses of the upper and lower rotors and the static blade pitch of the upper and lower rotors, the blade pitch of the upper and lower rotors is calculated (such as Figure 6 (indicated).

[0106] The key points of the present invention are:

[0107] (1) Establish the mathematical relationship between blade section bending moment load-structural strain-displacement response;

[0108] (2) Select a suitable blade profile group and calculate the blade displacement / deformation based on the measured load information of the selected profile;

[0109] To summarize, the present application provides a real-time calculation method for the propeller pitch of a coaxial twin-rotor, which monitors and measures the propeller pitch in real time during flight through the existing measured blade load information on the aircraft, realizes early warning of insufficient propeller pitch and provides reference information for flight control / manipulation to ensure flight safety.

[0110] Figure 1 In the figure: Z is the vertical coordinate (flapping direction); X is the spanwise coordinate; t is the time coordinate, corresponding to the azimuth angle; R is the rotor radius; FAero is the distributed aerodynamic force. Figure 2 The figure shows the distribution of the bending moment load at two sections of the blade as a function of the blade azimuth angle. Figure 3 Figure 2 shows the calculated results of blade displacement under actual blade loads during a flight. Figure 4 Figure 2 shows the calculated results of the blade tip pitch in a fixed coordinate system under the actual blade load measured during a flight.

Claims

1. A real-time calculation method for the pitch of coaxial twin rotors, characterized in that: The method comprises: Step 1: Using the analytical calculation information from the design phase, establish the mathematical relationship between the blade section bending moment load-structural strain-displacement response; Step 2: Select a suitable blade profile and solve the displacement response at the blade tip R based on the flight measured load data and the mathematical relationship between the blade profile bending moment load-structural strain-displacement response: Step 3: According to the flight measured load data M(r, t) of the blade profiles of the upper and lower rotors, the displacement responses of the upper and lower rotors are calculated respectively; according to the displacement responses of the upper and lower rotors and the static blade pitches of the upper and lower rotors, the blade pitches of the upper and lower rotors are calculated.

2. The method according to claim 1, characterized in that Step 1 includes: Step 11: Establish the flapping direction displacement response Z(x, t) of the blade at time t and spanwise position x: In the above formula, γ n is the vibration mode of each order of flapping mode, q n is the corresponding generalized modal displacement; Step 12: According to the flapping direction displacement response Z(x, t), the blade section bending moment load M(x, t) at time t and spanwise position x is obtained: Among them, EI flap (x) is the flapping stiffness, γ″ n (x) is the structural strain.

3. The method according to claim 1, characterized in that Step 2 includes: Step 21: Select the flight measured load data M(r, t) of the appropriate blade profile r: M(r,t)=Mb1(r)·q1(t)+Mb2(r)·q2(t)+Mb3(r)·q3(t)+…; Step 22: Calculate the displacement response Z(R, t) at the blade tip R: Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t)+…; Step 23: Obtain the first-order generalized modal displacement q1(t) through the flight measured load data M(r, t): Step 24: Substitute q1(t) into the displacement response Z(R, t) to obtain the displacement response at the blade tip R; 4. The method according to claim 1, characterized in that: Step 2 includes: Step 21: Select the flight measured load data of three blade sections with spanwise position coordinates a, b, and c: M(a, t), M(b, t), and M(c, t): M(a,t)=Mb1(a)·q1(t)+Mb2(a)·q2(t)+Mb3(a)·q3(t); M(b,t)=Mb1(b)·q1(t)+Mb2(b)·q2(t)+Mb3(b)·q3(t); M(c,t)=Mb1(c)·q1(t)+Mb2(c)·q2(t)+Mb3(c)·q3(t); Step 22: Solve the equations of M(a, t), M(b, t) and M(c, t) to obtain q1(t), q2(t) and q3(t); Step 23: Perform a third-order truncation on the displacement response at the blade tip R, and substitute q1(t), q2(t), and q3(t) to obtain the blade tip displacement Z(R, t): Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ2(R)·q3(t).

5. A real-time calculation device for the pitch of coaxial twin rotors, characterized in that: The device comprises a mathematical relationship establishing unit, a displacement response calculating unit and a propeller pitch calculating unit, wherein: A mathematical relationship building unit is used to establish the mathematical relationship between the blade section bending moment load-structural strain-displacement response using the analytical calculation information from the design phase; The displacement response calculation unit is used to select a suitable blade profile and solve the displacement response at the blade tip R according to the flight measured load data and the mathematical relationship between the blade profile bending moment load-structural strain-displacement response; The blade pitch calculation unit is used to calculate the displacement response of the upper and lower rotors according to the flight measured load data M(r, t) of the blade profiles of the upper and lower rotors; and calculate the blade pitch of the upper and lower rotors according to the displacement response of the upper and lower rotors and the static blade pitch of the upper and lower rotors.

6. The device according to claim 5, characterized in that The mathematical relationship establishment unit is specifically used for: Establish the flapping direction displacement response Z(x, t) of the blade at time t and spanwise position x: In the above formula, γ n is the vibration mode of each order of flapping mode, q n is the corresponding generalized modal displacement; According to the flapping direction displacement response Z(x, t), the blade section bending moment load M(x, t) at time t and spanwise position x is obtained: Among them, EI flap (x) is the flapping stiffness, γ″ n (x) is the structural strain.

7. The device according to claim 5, characterized in that The displacement response calculation unit is specifically used for: Select the flight measured load data M(r, t) of the appropriate blade profile r: M(r,t)=Mb1(r)·q1(t)+Mb2(r)·q2(t)+Mb3(r)·q3(t)+…; Calculate the displacement response Z(R, t) at the blade tip R: Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t)+…; The first-order generalized modal displacement q1(t) is obtained through the flight measured load data M(r, t): Substitute q1(t) into the displacement response Z(R, t) to obtain the displacement response at the blade tip R; 8. The device according to claim 5, characterized in that The displacement response calculation unit is specifically used for: The flight measured load data of three blade sections with spanwise position coordinates a, b, and c are selected: M(a, t), M(b, t), and M(c, t): M(a,t)=Mb1(a)·q1(t)+Mb2(a)·q2(t)+Mb3(a)·q3(t); M(b,t)=Mb1(b)·q1(t)+Mb2(b)·q2(t)+Mb3(b)·q3(t); M(c,t)=Mb1(c)·q1(t)+Mb2(c)·q2(t)+Mb3(c)·q3(t); Solve the system of equations M(a, t), M(b, t) and M(c, t) to obtain q1(t), q2(t) and q3(t); The displacement response at the blade tip R is truncated to the third order, and q1(t), q2(t) and q3(t) are substituted to obtain the blade tip displacement Z(R, t): Z(R,t)=γ1(R)·q1(t)+γ2(R)·q2(t)+γ3(R)·q3(t).

Images