A method for flight load trim of a twin intake aircraft

Abstract

The application discloses a two-side air intake type airplane flight load balancing method, which comprises the following steps: step one, determining airplane dynamic parameters obtained by solving a total maneuvering load; step two, calculating a wing surface type damping load and superimposing the damping load on original distributed load to obtain distributed load after superimposition of damping; step three, taking the distributed load after superimposition of damping as input balancing total force; step four, balancing a longitudinal moment; step five, balancing a lateral moment; and step six, balancing a side moment; the method directly adopts airplane dynamic parameters obtained by solving an initial airplane dynamics equation, comprehensively considers the aerodynamic load characteristics of various components, adds the influence amount of damping load, and effectively improves the precision of flight load data after balancing.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamic load design for aircraft, specifically, it relates to a method for balancing flight loads of aircraft with two-sided air intakes. Background Technology

[0002] In flight load calculations, the total load center of gravity and dynamic parameters obtained by solving the flight dynamics equations differ slightly from those obtained by calculating the total load center of gravity using distributed loads. Furthermore, the dynamic parameters obtained by recalculating the aircraft's dynamic parameters using the flight dynamics equations based on the total load center of gravity calculated using distributed loads also differ slightly. This difference arises because the total force and load obtained by integrating the distributed loads through theory or pressure testing differ from the focal position or aerodynamic coefficients used in the initial flight dynamics equations. This difference leads to overall aircraft imbalance. In flight load design, appropriate methods must be employed to correct this imbalance. Traditional methods for correcting overall aircraft imbalance mainly include two approaches: first, eliminating the calculation error of the aerodynamic center of gravity using the overall aircraft inertial force; second, applying corrective loads to the wing-fuselage interference area and the tail-fuselage interference area. While these two methods can eliminate overall aircraft imbalance in engineering, they still present the following problems: first, the influence of damping loads is not considered; and second, the error in the superimposed correction amount when correcting lateral moments is relatively large. Summary of the Invention

[0003] In view of the above-mentioned prior art, the purpose of this invention is to overcome the shortcomings of the prior art, adapt to the actual needs, and thus provide a method for balancing the flight load of a two-sided air intake aircraft to correct the overall load imbalance in flight load design.

[0004] To achieve the above objectives, the technical solution adopted by the present invention is: a method for flight load trimming of a two-sided air intake aircraft, the method comprising the following steps:

[0005] Step 1: Determine the aircraft dynamic parameters obtained when solving for the total maneuver load: ω x ω y ω z n y and n z ;

[0006] Step 2: Calculate the airfoil-like damping load using existing traditional engineering methods. The calculation method for airfoil-like damping loads is achieved through the following formula:

[0007]

[0008] μ0=l 2 / S (2)

[0009]

[0010]

[0011]

[0012]

[0013]

[0014] In formulas (1) to (7), the symbols have the following meanings:

[0015] S—Total wing surface area;

[0016] l — wingspan;

[0017] z — the z-coordinate of the face center point at the calculated mesh;

[0018] V – True speed of flight;

[0019] S(z) – The surface area of ​​the wing from the z-section to the wingtip;

[0020] ΔS — Calculate the grid area;

[0021] ω x —Roll angular velocity;

[0022] Δp zn —Calculate the damping load at the grid;

[0023] The above-mentioned wing-type damping load is superimposed on the original distributed load to obtain the superimposed damped distributed load.

[0024] Step 3: The distributed load after superimposed damping is used as the input total balancing force. The total balancing force is achieved through formulas (8) and (9), which are as follows:

[0025]

[0026]

[0027] Step 4: Balance the longitudinal moment:

[0028] The overall longitudinal torque imbalance is:

[0029]

[0030] The longitudinal moment imbalance is corrected by superimposing a load distribution on the fuselage. For conventional layout aircraft with side air intakes, the specific superimposed fuselage load distribution is as follows: a triangular distribution is taken in the forward L1 section of the fuselage, and a rectangular distribution is taken in the rear L0 section, as shown below.Figure 2 As shown;

[0031] Step 5: Balance the lateral moment:

[0032] The total lateral torque imbalance of the machine is:

[0033]

[0034] The lateral moment imbalance is corrected by coordinating the loads on the left and right wings. The specific correction formula is as follows:

[0035] F y_jy_all =F y_zjy +F y_yjy (12)

[0036] dF jy ·(Z yjy -Z zjy )=dM x (13)

[0037] F y_zjy_new =F y_zjy -dF jy (14)

[0038] F y_yjy_new =F y_zjy -dF jy (15);

[0039] Step 6: Balance the lateral moment:

[0040] The total lateral torque imbalance of the machine is:

[0041]

[0042] The lateral moment imbalance is corrected by superimposing a load distribution on the fuselage. For aircraft with conventional side-intake layouts, the specific load distribution pattern on the fuselage is as follows: a triangular distribution is taken in the forward L1 section of the fuselage, and a rectangular distribution is taken in the rear L0 section. That is, the load distribution pattern of the fuselage is consistent with the superimposed fuselage load distribution pattern in step four, thus completing the balance of the flight load for aircraft with side-intake layouts. Figure 2 As shown;

[0043] In formulas (8) to (16), the symbols have the following meanings:

[0044] —Roll acceleration;

[0045] —Yaw angular acceleration;

[0046] —Pitch acceleration;

[0047] ω x —Roll angular velocity;

[0048] ω y —Yaw angular velocity;

[0049] ω z —Pitch angular velocity;

[0050] n y —Longitudinal overload;

[0051] n z — Lateral overload;

[0052] G – Total weight of the aircraft;

[0053] I xx —Moment of inertia about the X-axis;

[0054] I yy —Moment of inertia about the Y-axis;

[0055] I zz —Moment of inertia about the Z-axis;

[0056] I xy —The product of inertia about the X and Y axes;

[0057] m — the total number of components involved in balancing;

[0058] F yi —Total normal load of all components;

[0059] F zi —Total lateral load of all components;

[0060] X i —X-axis centering of each component;

[0061] Y i —Y-axis pressure center of each component;

[0062] Z i —Z-axis pressure center of each component;

[0063] F y_zjy —Longitudinal load on the left wing;

[0064] F y_yjy —Longitudinal load on the right wing;

[0065] Z zjy —Z-axis coordinate of the spanwise center of pressure of the left wing;

[0066] Z yjy —Z-axis coordinate of the right wing spanwise center of pressure;

[0067] Fy_zjy_new —Coordinate the longitudinal load on the left wing;

[0068] F y_yjy_new —Coordinate the longitudinal load on the right wing.

[0069] Furthermore, the coordinate system used in steps one to six of this invention is specifically as follows: the origin is at the center of gravity, the x-axis is parallel to the horizontal axis of the aircraft pointing forward, the y-axis points upward within the plane of symmetry of the aircraft, and the z-axis is determined according to the right-hand rule, as shown below. Figure 1 As shown.

[0070] Furthermore, the values ​​of L1 and L0 are determined by the specific aerodynamic characteristics of the aircraft fuselage.

[0071] This invention directly uses the aircraft dynamic parameters obtained during the solution of the total maneuvering load; it calculates the damping load of the wing-type distributed load data, and then superimposes them to obtain distributed load data considering the damping effect; the total force is corrected using the coefficient method, and the longitudinal moment adopts a scheme to correct the fuselage aerodynamic load distribution. For conventional layout aircraft with side air intakes, the superimposed fuselage load distribution is as follows: Figure 2 The distribution pattern shown is as follows: a triangular distribution is taken in the front L1 part of the fuselage, and a rectangular distribution is taken in the rear L0 part of the fuselage. The values ​​of L1 and L0 are determined by the specific aerodynamic characteristics of the aircraft fuselage. The lateral moment adopts a scheme that coordinates the loads of the left and right wings. Finally, the lateral moment is corrected by adopting the same scheme as the longitudinal moment correction mentioned above, and the overall unbalance is finally corrected.

[0072] The beneficial effects of the present invention are as follows: The method of the present invention directly uses the aircraft dynamic parameters obtained by solving the initial aircraft dynamic equations, comprehensively considers the aerodynamic load characteristics of each component, and adds the influence of damping load, which effectively improves the accuracy of the flight load data after trimming. Attached Figure Description

[0073] Figure 1 This is a schematic diagram of the coordinate system of the present invention, with the origin being the centroid;

[0074] Figure 2 This invention includes an additional fuselage load distribution diagram;

[0075] Figure 3 This is a flowchart illustrating a specific implementation of the present invention.

[0076] Where L1 is the length of the triangular distribution area, L0 is the length of the rectangular distribution area, and ω x X is the roll angular velocity. i For each component, the X-axis pressure center, Y-axis pressure center i For the Y-axis pressure center of each component, Z i For the Z-axis pressure center of each component, ω y Let ω be the angular velocity of the heading. zM is the pitch angular velocity. x M is the rolling torque. y For the yaw moment, M z For pitching moment, δ x For the aileron deflection angle, δ y For rudder deflection, δ z This refers to the elevator deflection angle. Detailed Implementation

[0077] The specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are for illustration and explanation only and are not intended to limit the present invention.

[0078] Example

[0079] like Figures 1 to 3 As shown, this invention provides a method for flight load trim of a two-sided air intake aircraft. This method directly uses the aircraft dynamic parameters obtained during the calculation of the total maneuver load to calculate the damping load of the wing-type distributed load data. The coordinate system used in this invention is as follows: Figure 1 As shown, the coordinate system is as follows: the origin is at the center of gravity, the x-axis is parallel to the horizontal axis of the aircraft pointing forward, the y-axis points upward in the plane of symmetry of the aircraft, and the z-axis is determined according to the right-hand rule. This method includes the following steps:

[0080] Step 1: Determine the aircraft dynamic parameters obtained when solving for the total maneuver load: ω x ω y ω z n y and n z ;

[0081] Step 2: Calculate the airfoil-like damping load using the following method. The specific calculation method for the airfoil-like damping load is as follows:

[0082]

[0083] μ0=l 2 / S (2)

[0084]

[0085]

[0086]

[0087]

[0088]

[0089] In formulas (1) to (7), the meanings of each symbol are as follows:

[0090] s—total wing surface area;

[0091] l — wingspan;

[0092] z — the z-coordinate of the face center point at the calculated mesh;

[0093] V – True speed of flight;

[0094] S(z) – The surface area of ​​the wing from the z-section to the wingtip;

[0095] ΔS — Calculate the grid area;

[0096] ω x —Roll angular velocity;

[0097] Δp zn —Calculate the damping load at the grid;

[0098] The above-mentioned wing-type damping load is superimposed on the original distributed load to obtain the superimposed damped distributed load.

[0099] Step 3: Using the superimposed damped distributed load as input, balance the total force according to the following formula:

[0100]

[0101]

[0102] Step 4: Balance the longitudinal moment:

[0103] The overall longitudinal torque imbalance is:

[0104]

[0105] The longitudinal moment imbalance is corrected by superimposing a load distribution on the fuselage. For conventionally configured aircraft with side air intakes, the superimposed fuselage load distribution is taken as an example. Figure 2 The distribution pattern shown is as follows: a triangular distribution is taken in the front L1 part of the fuselage, and a rectangular distribution is taken in the rear L0 part of the fuselage. The values ​​of L1 and L0 are determined by the specific aerodynamic characteristics of the aircraft fuselage.

[0106] Step 5: Balance the lateral moment:

[0107] The total lateral torque imbalance of the machine is:

[0108]

[0109] The lateral moment imbalance is corrected by coordinating the loads on the left and right wings. The specific correction formula is as follows:

[0110] F y_jy_all =F y_zjy +F y_yjy (12)

[0111] dF jy ·(Z yjy -Z zjy )=dM x (13)

[0112] F y_yjy_new =F y_zjy -dF jy (14)

[0113] F y_yjy_new =F y_zjy -dF jy (15);

[0114] Step 6: Balance the lateral moment:

[0115] The total lateral torque imbalance of the machine is:

[0116]

[0117] The lateral moment imbalance is corrected by superimposing a load distribution on the fuselage. For conventionally configured aircraft with side air intakes, the fuselage load distribution adopts the same distribution pattern as the superimposed fuselage load distribution in step four. Specifically, the distribution is triangular in the forward L1 section of the fuselage and rectangular in the rear L0 section. The values ​​of L1 and L0 are determined by the specific aerodynamic characteristics of the aircraft fuselage. Figure 2 As shown;

[0118] In formulas (8) to (16), the symbols have the following meanings:

[0119] —Roll acceleration;

[0120] —Yaw angular acceleration;

[0121] —Pitch acceleration;

[0122] ω x —Roll angular velocity;

[0123] ω y —Yaw angular velocity;

[0124] ω z —Pitch angular velocity;

[0125] n y —Longitudinal overload;

[0126] n z — Lateral overload;

[0127] G – Total weight of the aircraft;

[0128] I xx —Moment of inertia about the X-axis;

[0129] I yy —Moment of inertia about the Y-axis;

[0130] I zz —Moment of inertia about the Z-axis;

[0131] I xy —The product of inertia about the X and Y axes;

[0132] m — the total number of components involved in balancing;

[0133] F yi —Total normal load of all components;

[0134] F zi —Total lateral load of all components;

[0135] X i —X-axis centering of each component;

[0136] Y i —Y-axis pressure center of each component;

[0137] Z i —Z-axis pressure center of each component;

[0138] F y_zyj —Longitudinal load on the left wing;

[0139] F y_yjy —Longitudinal load on the right wing;

[0140] Z zjy —Z-axis coordinate of the spanwise center of pressure of the left wing;

[0141] Z yjy —Z-axis coordinate of the right wing spanwise center of pressure;

[0142] F y_zjy_new —Coordinate the longitudinal load on the left wing;

[0143] F y_yjy_new —Coordinate the longitudinal load on the right wing.

[0144] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

[0145] Furthermore, those skilled in the art will understand that although some embodiments herein include certain features included in other embodiments but not others, combinations of features from different embodiments are meant to be within the scope of the invention and form different embodiments. For example, in the foregoing claims, any of the claimed embodiments can be used in any combination. The information disclosed in this background section is intended only to enhance the understanding of the general background of the invention and should not be construed as an admission or in any way implying that such information constitutes prior art known to those skilled in the art.

Claims

1. A type of aircraft with side air intakes for flight The load balancing method is characterized by: The method includes the following steps: Step 1: Determine the aircraft dynamic parameters obtained when solving for the total maneuver load: ω x ω y ω z n y and n z ; Step 2: Calculate the airfoil-like damping load using existing traditional engineering methods. The calculation method for airfoil-like damping loads is achieved through the following formula: μ0=l 2 / S (2) In formulas (1) to (7), the symbols have the following meanings: S—Total wing surface area; l — wingspan; z — the z-coordinate of the face center point at the calculated mesh; V – True speed of flight; S(z) – The surface area of ​​the wing from the z-section to the wingtip; ΔS — Calculate the grid area; ω x —Roll angular velocity; Δp zn —Calculate the damping load at the grid; The above-mentioned wing-type damping load is superimposed on the original distributed load to obtain the superimposed damped distributed load. Step 3: The distributed load after superimposed damping is used as the input total balancing force. The total balancing force is achieved through formulas (8) and (9), which are as follows: Step 4: Balance the longitudinal moment: The overall longitudinal torque imbalance is: The longitudinal moment imbalance is corrected by superimposing load distribution on the fuselage. For conventional layout aircraft with side air intakes, the superimposed fuselage load distribution is as follows: a triangular distribution is taken in the front L1 part of the fuselage, and a rectangular distribution is taken in the rear L0 part of the fuselage. Step 5: Balance the lateral moment: The total lateral torque imbalance of the machine is: The lateral moment imbalance is corrected by coordinating the loads on the left and right wings. The specific correction formula is as follows: F y_jy_all =F y_zjy +F y_yjy (12) dF jy ·(Z yjy -Z zjy )=dM x (13) F y_zjy_new =F y_zjy -dF jy (14) F y_yjy_new =F y_zjy -dF jy (15); Step 6: Balance the lateral moment: The total lateral torque imbalance of the machine is: The lateral moment imbalance is corrected by superimposing load distribution on the fuselage. For conventional layout aircraft with two air intakes, the load distribution pattern of the fuselage is taken to be consistent with the fuselage load distribution pattern superimposed in step four, thus completing the balance of the flight load of the aircraft with two air intakes. In formulas (8) to (16), the symbols have the following meanings: —Roll acceleration; —Yaw angular acceleration; —Pitch acceleration; ω x —Roll angular velocity; ω y —Yaw angular velocity; ω z —Pitch angular velocity; n y —Longitudinal overload; n z — Lateral overload; G – Total weight of the aircraft; I xx —Moment of inertia about the X-axis; I yy —Moment of inertia about the Y-axis; I zz —Moment of inertia about the Z-axis; I xy —The product of inertia about the X and Y axes; m — the total number of components involved in balancing; F yi —Total normal load of all components; F zi —Total lateral load of all components; X i —X-axis centering of each component; Y i —Y-axis pressure center of each component; Z i —Z-axis pressure center of each component; F y_zjy —Longitudinal load on the left wing; F y_yjy —Longitudinal load on the right wing; Z zjy —Z-axis coordinate of the spanwise center of pressure of the left wing; Z yjy —Z-axis coordinate of the right wing spanwise center of pressure; F y_zjy_new —Coordinate the longitudinal load on the left wing; F y_yjy_new —Coordinate the longitudinal load on the right wing.

2. The method for trimming flight loads of a two-sided air intake aircraft according to claim 1, characterized in that: The coordinate system used in steps one through six is ​​as follows: the origin is at the center of gravity, the x-axis is parallel to the horizontal axis of the aircraft and points forward, the y-axis points upward in the plane of symmetry of the aircraft, and the z-axis is determined according to the right-hand rule.

3. The method for balancing flight loads of a two-sided air intake aircraft according to claim 1, characterized in that: The values ​​of L1 and L0 are determined by the specific aerodynamic characteristics of the aircraft fuselage.

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