Method for calculating deflection angle of catapult-assisted take-off airframe in case of failure of unilateral engine

Abstract

The invention belongs to the technical field of shipboard aircraft catapult, and particularly relates to a method for calculating the deflection angle of a catapult-assisted take-off airframe in unilateral engine failure, which comprises the following steps of: respectively establishing a static balance equation of a catapult tensioning stage, a kinetic equation of a catapult taxiing stage and a kinetic equation of a free taxiing stage; detailed dynamic analysis is carried out on all stages of the catapult-assisted take-off process of the carrier-based aircraft with the engine failure on one side, and a corresponding dynamic differential equation set is established. And the change conditions of the course acceleration, the speed and the deflection angle of the aircraft along with time can be theoretically calculated. According to the provided theoretical method, factors are comprehensively considered, the calculation speed is high, implementation is easy, relevant parameters can be replaced for calculation according to actual conditions during application, and theoretical support can be provided for rapid evaluation of the feasibility of catapult-assisted take-off of a shipboard aircraft with a wing-hung engine under the condition that the engine on one side fails and aircraft attitude prediction.

Description

Technical Field

[0001] The present application belongs to the technical field of carrier-based aircraft catapults, and in particular relates to a method for calculating the deflection angle of a launch vehicle body during a catapult launch when a single-side engine fails. Background Art

[0002] During the catapult takeoff process, carrier-based aircraft must withstand loads such as engine thrust, catapult force, aircraft lift, air resistance, friction between the main landing gear tires and the ship surface, and inertia. Under normal circumstances, the aircraft load is basically in a left-right symmetrical state. At present, the dynamic models for the catapult takeoff of carrier-based aircraft are all established under the condition that the aircraft itself is in a normal state. For example, in the model established in reference [1], the aircraft position and load are symmetrical along the heading, and all directions except the heading are in a force balance state, which is beneficial for analyzing only the heading overload of the aircraft. In reference [2], it is considered that due to the asymmetry of the aircraft position, the aircraft symmetry plane and the catapult track cannot be aligned, resulting in an initial eccentricity before ejection. The initial eccentricity makes the aircraft load condition more complicated and affects the aircraft's yaw motion, but the aircraft itself is in a normal state.

[0003] The abnormal state of the carrier-based aircraft's own engine may also cause the aircraft to undergo complex attitude changes during the catapult process, which is also a type of working condition that deserves special consideration for the catapult of carrier-based aircraft. For carrier-based aircraft with wing-mounted engines, when one engine fails, the other engine will not only generate heading thrust, but also considerable turning torque, putting the aircraft in an asymmetric stress state. After searching, no theoretical research work on the dynamic analysis of catapult takeoff of carrier-based aircraft in this state has been found.

[0004] Therefore, how to comprehensively consider and theoretically calculate the changes of the deflection angle and angular velocity of carrier-based aircraft under multiple loads over time is a problem that needs to be solved. Summary of the invention

[0005] The purpose of this application is to provide a method for calculating the deflection angle of a catapult launch body when a single-side engine fails, so as to solve the problem that there is no existing theory for dynamic analysis of catapult launch of carrier-based aircraft under asymmetric stress state.

[0006] The technical solution of the present application is: a method for calculating the deflection angle of a body launched by ejection when one side of the engine fails, comprising:

[0007] During the catapult launch phase, a global coordinate system is established based on the aircraft's current heading, and the calculation conditions and necessary parameters for the catapult launch of the aircraft are determined by assuming that one engine fails and the other engine operates normally.

[0008] In the ejection tensioning stage, all the forces on the aircraft in the current state are obtained, and all the forces are split into the resultant force in the X-axis direction, the resultant force in the Z-axis direction and the pitching moment at the center of gravity, and the static equilibrium equations of the ejection tensioning stage are established in different directions;

[0009] During the ejection roll phase, obtain the rolling friction f between the aircraft's front and main landing gear tires and the ship's surface n and f m , lift L, heading resistance D, assuming that the lift and heading resistance of the aircraft act on the center of gravity of the aircraft, perform force analysis on the xoz plane in the global coordinate system, and establish the dynamic equation of the ejection run phase based on the force analysis results;

[0010] In the free sliding stage, the ejection force T is cancelled, and the force analysis is performed in the X-axis direction, Z-axis direction and center of gravity position respectively, and the dynamic equation of the free sliding stage is established by combining the static equilibrium equation of the ejection tensioning stage and the dynamic equation of the ejection sliding stage;

[0011] Substitute the corresponding initial conditions of each stage, solve the dynamic equations of the ejection run stage and the free run stage, and output the change of the body deflection angle data over time.

[0012] Preferably, the necessary parameters include:

[0013] 1) Aircraft mass m, unit: kg;

[0014] 2) Aircraft center of gravity coordinates (x) m ,y m ,z m ), unit: m;

[0015] 3) The moment of inertia of the aircraft around the center of gravity z axis J z , unit: kg·m 2 ;

[0016] 4) Wing equivalent area S, unit: m 2 ;

[0017] 5) Aircraft lift coefficient C L , dimensionless;

[0018] 6) Aircraft drag coefficient C D , dimensionless;

[0019] 7) Engine action point coordinates (x e ,y e ,z e ), unit: m;

[0020] 8) Coordinates of the front landing gear tire contact point (x n ,y n ,zn ), unit: m;

[0021] 9) Coordinates of the left main landing gear tire contact point (x ml ,y ml ,z ml ), unit: m;

[0022] 10) Coordinates of the right main landing gear tire contact point (x mr ,y mr ,z mr ), unit: m;

[0023] 11) Angle α between the engine thrust line and the horizontal direction, unit: °;

[0024] 12) Angle β between the catapult rod and the ship deck, unit: °;

[0025] 13) Angle γ between the restraining rod and the ship deck, unit: °;

[0026] 14) The height of the connection point between the ejection rod and the restraining rod and the front landing gear from the ship surface, unit: m;

[0027] 15) Engine thrust P, unit: N;

[0028] 16) Atmospheric density ρ, unit: kg·m -3 ;

[0029] 17) The rolling friction coefficient between the front landing gear tire and the ship deck is μ1, dimensionless;

[0030] 18) The rolling friction coefficient between the main landing gear tire and the ship deck is μ2, dimensionless;

[0031] 19) The lateral sliding friction coefficient between the main landing gear tire and the ship deck is μ3, dimensionless;

[0032] 20) The relationship between the ejection force T and the stroke l provided by the catapult.

[0033] Preferably, in the global coordinate system, the aircraft's forward direction is taken as the negative direction of the x-axis, the z-axis is perpendicular to the ship's surface and upward as the positive direction, and the y-axis is perpendicular to the xoz plane and points to the right side of the aircraft.

[0034] Preferably, when determining the calculation conditions during the catapult takeoff phase of the carrier-based aircraft, the lateral acceleration and pitch acceleration of the aircraft are ignored, the influence of the swaying of the aircraft carrier on the sea surface and the unevenness of the ship surface on the aircraft load is ignored, and the influence of atmospheric disturbances on the aircraft load is ignored.

[0035] Preferably, when determining the calculation conditions during the catapult takeoff phase of the carrier-based aircraft, it is assumed that the fuselage and the landing gear are both rigid bodies, tire compression is not considered, and the center of gravity, the front landing gear, the ejection rod and the check rod are located on the symmetry plane of the aircraft.

[0036] Preferably, during the ejection tensioning stage, the forces acting on the aircraft include: the thrust P of the right engine, the ejection force T of the ejection rod, and the tension F of the restraining rod. h , the ground reaction force N1 borne by the front landing gear, the ground reaction force N2 borne by the single main landing gear, and the gravity mg.

[0037] Preferably, in the catapult tensioning stage, according to the resultant force of the aircraft in the x-axis direction being 0, the static equilibrium equation is established as:

[0038] ΣF x = -P cos α-T cos β-F h cosγ=0

[0039] According to the resultant force of the aircraft in the z-axis direction is 0, the static equilibrium equation is established:

[0040] ∑F z =N1+2N2+P sinα-T sinβ-F h sinγ-mg=0

[0041] According to the pitch moment of the aircraft at the center of gravity is 0, we can get:

[0042] ∑M z =N1(x m -x n )-2N2(x ml -x m )+T[(-z n -h+z m )cosβ-(x m -x n )sinβ]-F h [(-z n -h+z m )cosγ-(x m -x n )sinγ]-P[(z e -z m )cosα+(x e -x m )sinα]=0.

[0043] Preferably, during the ejection roll phase, it is assumed that the aircraft body deflects around the front main landing gear due to the torque imbalance under the thrust of the right engine. During the deflection of the aircraft body, the main landing gear tires are subjected to a sliding friction force f with the ship surface. c ;but:

[0044] According to the resultant force in the z-axis direction, the equilibrium equation is established:

[0045] ΣF z=N1+2N2+P sinα-T sinβ-mg+L=0

[0046] The dynamic equation is established based on the resultant force in the x-axis direction:

[0047]

[0048] Preferably, the rolling friction force f n 、f m , sliding friction f c And the formula for determining the aircraft lift L and aircraft drag D is:

[0049] f n =μ1N1

[0050] f m =μ2N2

[0051] f c =μ3N2

[0052]

[0053] Where v is the aircraft heading speed.

[0054] Preferably, when the aircraft deflects around the front landing gear, the dynamic equation for the deflection of the aircraft fuselage is established as:

[0055]

[0056] in is the total moment of the aircraft about the front wheel, is the aircraft's moment of inertia about the front wheel.

[0057] Preferably, the dynamic equation of the ejection run phase is:

[0058]

[0059] Preferably, the dynamic equation of the free sliding stage is:

[0060] N1+2N2+P sinα-mg+L=0

[0061]

[0062] J z θ(t)″=y e P-2f c (x ml -x m ).

[0063] Preferably, the dynamic equations of the ejection run phase and the dynamic equations of the free run phase are numerically solved by using the ode45 function in MATLAB.

[0064] Preferably, the body deflection angle data includes heading displacement x, heading speed x′, deflection angle θ and deflection angular speed θ′.

[0065] The method for calculating the deflection angle of the body during catapult launch when one side of the engine fails in the present application performs a detailed dynamic analysis of each stage of the catapult launch process of a carrier-based aircraft with one side of the engine failing, and establishes a corresponding set of dynamic differential equations. The aircraft's heading acceleration, velocity, and deflection angle changes over time can be theoretically calculated. The provided theoretical method takes into account comprehensive factors, has a fast calculation speed, and is easy to implement. When applied, relevant parameters can be replaced for calculation according to actual conditions. It can quickly evaluate and provide theoretical support for the feasibility of catapult launch of a carrier-based aircraft with a wing-mounted engine when one side of the engine fails, as well as the aircraft attitude prediction. BRIEF DESCRIPTION OF THE DRAWINGS

[0066] In order to more clearly illustrate the technical solution provided by the present application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of the present application.

[0067] Figure 1 This is a schematic diagram of the overall process of this application;

[0068] Figure 2 This is the force analysis diagram of the carrier-based aircraft catapult tensioning stage in this application;

[0069] Figure 3 This is the force analysis diagram of the xoz plane during the catapult-running phase of the carrier-based aircraft in this application;

[0070] Figure 4 This is the force analysis diagram in the xoy plane during the catapult-running phase of the carrier-based aircraft in this application;

[0071] Figure 5 This is the curve of ejection force T and stroke l in this application;

[0072] Figure 6 This is the curve of heading displacement and speed changing with time for this application;

[0073] Figure 7 This is the curve of the change of the fuselage deflection angle over time for this application;

[0074] Figure 8 This is the curve of the fuselage deflection angular velocity changing with time in this application. DETAILED DESCRIPTION

[0075] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0076] A method for calculating the deflection angle of a catapult launch vehicle when one side of the engine fails, such as Figure 1 As shown, the following steps are included:

[0077] Step S100, during the catapult launch phase of the carrier-based aircraft, a global coordinate system is established according to the current forward direction of the aircraft, and the calculation conditions and necessary parameters for the catapult launch of the carrier-based aircraft are determined by assuming that one engine fails and the other engine operates normally.

[0078] The catapult takeoff process of carrier-based aircraft can generally be divided into three stages: catapult tensioning stage, catapult run stage and free run stage. Assume that the fuselage and landing gear are rigid bodies, tire compression is not considered, and the center of gravity, front landing gear, catapult rod and restraining rod are located on the symmetry plane of the aircraft. Establish a global coordinate system, take the aircraft's forward direction as the negative direction of the x-axis, the z-axis perpendicular to the ship's surface upward as the positive direction, and the y-axis perpendicular to the xoz plane pointing to the right side of the aircraft. Assume that one engine fails and the other engine operates normally; for example, the left engine of the aircraft fails and the right engine operates normally. During the catapult process, ignore the aircraft's lateral acceleration and pitch acceleration, ignore the impact of the carrier's swaying on the sea surface and the unevenness of the ship's surface on the aircraft load, and ignore the impact of atmospheric disturbances on the aircraft load.

[0079] After determining the calculation conditions, the following necessary parameters must be determined:

[0080] 1) Aircraft mass m, unit: kg;

[0081] 2) Aircraft center of gravity coordinates (x) m ,y m ,z m ), unit: m;

[0082] 3) The moment of inertia of the aircraft around the center of gravity z axis J z , unit: kg·m 2 ;

[0083] 4) Wing equivalent area S, unit: m 2 ;

[0084] 5) Aircraft lift coefficient C L , dimensionless;

[0085] 6) Aircraft drag coefficient C D , dimensionless;

[0086] 7) Engine action point coordinates (x e ,y e ,z e ), unit: m;

[0087] 8) Coordinates of the front landing gear tire contact point (x n ,y n ,z n ), unit: m;

[0088] 9) Coordinates of the left main landing gear tire contact point (x ml ,y ml ,z ml ), unit: m;

[0089] 10) Coordinates of the right main landing gear tire contact point (x mr ,y mr ,z mr ), unit: m;

[0090] 11) Angle α between the engine thrust line and the horizontal direction, unit: °;

[0091] 12) Angle β between the catapult rod and the ship deck, unit: °;

[0092] 13) Angle γ between the restraining rod and the ship deck, unit: °;

[0093] 14) The height of the connection point between the ejection rod and the restraining rod and the front landing gear from the ship surface, unit: m;

[0094] 15) Engine thrust P, unit: N;

[0095] 16) Atmospheric density ρ, unit: kg·m -3 ;

[0096] 17) The rolling friction coefficient between the front landing gear tire and the ship deck is μ1, dimensionless;

[0097] 18) The rolling friction coefficient between the main landing gear tire and the ship deck is μ2, dimensionless;

[0098] 19) The lateral sliding friction coefficient between the main landing gear tire and the ship deck is μ3, dimensionless;

[0099] 20) The relationship between the ejection force T (unit: N) and the stroke l (unit: m) provided by the catapult.

[0100] Step S200, in the ejection tensioning stage, obtain all the forces on the aircraft in the current state, and split all the forces on the aircraft into the resultant force in the X-axis direction, the resultant force in the Z-axis direction and the pitching moment at the center of gravity, and establish the static equilibrium equations of the ejection tensioning stage in different directions.

[0101] During the catapult tensioning stage, the aircraft is in a static state. At this time, the catapult loads the ejection force to the ejection rod, the ejection rod is tensioned, and the aircraft is in a static equilibrium state. The forces it is subjected to include: the thrust P of the right engine, the ejection force T of the ejection rod, and the tension F of the restraining rod. h , the ground reaction force N1 borne by the front landing gear, the ground reaction force N2 borne by the single main landing gear, and the gravity mg. Figure 2 shown.

[0102] The aircraft is in a state of static equilibrium. According to the resultant force of the aircraft in the x-axis direction being 0, the static equilibrium equation is established:

[0103] ΣF x = -P cos α-T cos β-F h cosγ=0 (1)

[0104] According to the resultant force of the aircraft in the z-axis direction is 0, the static equilibrium equation is established:

[0105] ΣF z =N1+2N2+P sinα-T sinβ-F h sinγ-mg=0 (2)

[0106] According to the pitch moment of the aircraft at the center of gravity is 0, we can get:

[0107]

[0108] Step S300, during the ejection roll phase, obtain the rolling friction force f between the aircraft's front and main landing gear tires and the deck n and f m , lift L, and heading resistance D. Assuming that the lift and heading resistance of the aircraft act on the center of gravity of the aircraft, the force analysis of the xoz plane is carried out in the global coordinate system, and the dynamic equation of the ejection run phase is established based on the force analysis results.

[0109] At this stage, the aircraft no longer bears the tension of the restraining rod. Under the ejection force of the ejection rod and the thrust of the right engine, the aircraft rushes out quickly until the ejection force disappears at the end of the catapult stroke. During this process, the aircraft has a large heading overload and the aircraft speed increases rapidly. Compared with the catapult tensioning stage, the aircraft also needs to withstand the rolling friction between the front and main landing gear tires and the ship surface. n and f m , lift L, heading resistance D, and the tension of the check rod disappears. Assuming that the lift and heading resistance of the aircraft act on the center of gravity of the aircraft, the force analysis of the aircraft in the xoz plane at this stage is as follows: Figure 3 shown.

[0110] In addition, due to the unbalanced torque under the thrust of the right engine, the aircraft may deflect around the front main landing gear. During the deflection of the aircraft, the main landing gear tires must also withstand the sliding friction force f with the ship surface. c , the stress condition is as follows Figure 4 shown.

[0111] The equilibrium equation is established based on the resultant force in the z-axis direction:

[0112] ∑F z =N1+2N2+P sinα-T sinβ-mg+L=0 (4)

[0113] The dynamic equation is established based on the resultant force in the x-axis direction:

[0114]

[0115] The rolling friction force f n 、f m , sliding friction f c The aircraft lift L and aircraft drag D are determined by the following formula:

[0116] f n =μ1N1 (6)

[0117] f m =μ2N2 (7)

[0118] f c =μ3N2 (8)

[0119]

[0120] Where v is the aircraft heading speed.

[0121] Considering the deflection of the aircraft around the front landing gear, the dynamic equation of the aircraft fuselage deflection is established:

[0122]

[0123] in is the total moment of the aircraft about the front wheel, is the moment of inertia of the aircraft around the front wheel. According to the parallel axis theorem, substitute the moment of inertia at the center of gravity J z calculate

[0124]

[0125] The total moment of the aircraft about the front wheel is:

[0126]

[0127] Substituting formula (12) and (13) into formula (11), we obtain:

[0128]

[0129] Step S400, in the free running stage, the ejection force T is canceled, and force analysis is performed in the X-axis direction, the Z-axis direction and the center of gravity position respectively, and the dynamic equation of the free running stage is established by combining the static equilibrium equation of the ejection tensioning stage and the dynamic equation of the ejection running stage.

[0130] During the free roll phase, the catapult rod is separated from the carrier aircraft, and the ejection force on the aircraft disappears compared to the catapult roll phase. Figure 3 and Figure 4 In the force analysis, cancel the ejection force T, and the dynamic equation can be established as follows:

[0131] N1+2N2+P sinα-mg+L=0 (15)

[0132]

[0133] J z θ(t)″=y e P-2f c (x ml -x m ) (17).

[0134] Step S500, substitute the initial conditions corresponding to each stage, solve the dynamic equations of the ejection run stage and the free run stage, and output the change of the body deflection angle data over time. The body deflection angle data includes heading displacement x, heading speed x', deflection angle θ and deflection angle speed θ'.

[0135] For the static equilibrium state of the ejection tensioning stage, the landing gear load can be easily obtained by solving the equation group. In the ejection run and free run stages, the dynamic equations of each stage need to be solved. The differential equations in steps S300-S400 can be numerically solved by the ode45 function in MATLAB. Substituting the corresponding initial conditions of each stage, the aircraft's heading displacement x, heading velocity x', deflection angle θ and deflection angular velocity θ' can be calculated over time t during the ejection run and free run stages.

[0136] Through the above design, a detailed dynamic analysis is conducted on each stage of the catapult takeoff process of a carrier-based aircraft with one engine failure, and the corresponding dynamic differential equations are established. The aircraft's heading acceleration, velocity, and deflection angle can be theoretically calculated over time. The provided theoretical method takes into account comprehensive factors, has a fast calculation speed, and is easy to implement. When applied, the relevant parameters can be replaced according to the actual situation for calculation. It can quickly evaluate and provide theoretical support for the feasibility of catapult takeoff of a carrier-based aircraft with a wing-mounted engine when one engine fails and the aircraft attitude prediction.

[0137] As a specific implementation, a specific example is used to illustrate:

[0138] Since the ejection roll phase is the phase in which the aircraft is subjected to the most complex stress during the ejection process, the static calculation of the ejection tensioning phase in step S200 is skipped in this example, and the free roll phase in step S400 is not discussed again. The ejection roll phase in step S300 is directly subjected to stress analysis and numerical calculation is performed according to step S500.

[0139] The details are as follows:

[0140] Step 1: Determine the calculation conditions and necessary parameters.

[0141] Assume the following parameters:

[0142] 1) Aircraft mass m = 15000 kg;

[0143] 2) Aircraft center of gravity coordinates (7.5m, 0m, 1m);

[0144] 3) The moment of inertia of the aircraft around the center of gravity z axis J z =250000kg·m 2 ;

[0145] 4) Wing equivalent area S = 50m 2 ;

[0146] 5) Aircraft lift coefficient C L =1.2;

[0147] 6) Aircraft drag coefficient C D =0.1;

[0148] 7) Engine action point coordinates (7.6m, 3m, 1.5m);

[0149] 8) Coordinates of the front landing gear tire contact point (1m, 0m, -1.6m);

[0150] 9) Coordinates of the left main landing gear tire contact point (8m, -2.5m, -1.6m);

[0151] 10) Coordinates of the right main landing gear tire contact point (8m, 2.5m, -1.6m);

[0152] 11) The angle between the engine thrust line and the horizontal direction is α = 3°;

[0153] 12) The angle between the catapult rod and the ship deck is β = 45°;

[0154] 13) The angle between the restraining rod and the ship deck is γ = 45°;

[0155] 14) The height of the connection point between the ejection rod, the check rod and the front landing gear from the ship surface is h = 1.2m;

[0156] 15) Engine thrust P = 10000N;

[0157] 16) Atmospheric density ρ = 1.2254 kg·m -3 ;

[0158] 17) The rolling friction coefficient between the front landing gear tire and the ship deck is μ1 = 0.2;

[0159] 18) The rolling friction coefficient between the main landing gear tire and the ship deck is μ2 = 0.2;

[0160] 19) The lateral sliding friction coefficient between the main landing gear tire and the ship deck is μ3 = 0.4;

[0161] 20) The relationship between the ejection force T provided by the catapult and the stroke l is as follows Figure 5 As shown, when the stroke is 0, the catapult provides 3×10 5 N ejection force, which is kept balanced during the ejection tensioning stage. When the stroke is 80 m, the ejection rod separates from the front landing gear of the carrier aircraft, and the ejection force is 0. The ejection force variation curve can be expressed by a piecewise function, as shown in formula (18):

[0162]

[0163] Step 2: Establish the static equilibrium equation of the ejection tensioning stage.

[0164] slightly.

[0165] Step 3: Establish the dynamic equation of the ejection run phase.

[0166] According to the equilibrium equation of the ejection run phase, the expressions of the vertical ground loads N1 and N2 of the front and main landing gear can be derived:

[0167]

[0168] N1=mg+T sinβ-LP sinα-2N2 (20)

[0169] Step 4: Establish the dynamic equation of the free sliding stage.

[0170] slightly.

[0171] Step 5: Solve the dynamic equations and output the changes of deflection angle, angular velocity and other parameters over time.

[0172] Substitute formula (19) and (20) into formula (5) and (14), and numerically solve the differential equations by using the ode45 function in MATLAB. Substitute the parameters set in step 1, and set the initial conditions to [0, 0, π / 36, π / 18], that is, the initial heading displacement x0 = 0m, the initial heading speed x0′ = 0m / s, the initial deflection angle θ0 = 5°, and the initial deflection angular velocity θ0′ = 10° / s. The changes of the heading displacement, heading speed, fuselage deflection angle, and fuselage deflection angular velocity of the aircraft over time during the catapult roll can be obtained, as shown in Figure 6 to Figure 8 As shown. It can be seen from the figure that when the ejection displacement reaches 80m, the ejection force disappears, and the ejection run process ends at t=3.012s, and the aircraft's heading speed is 65.45m / s. Due to the existence of the initial angular velocity, the fuselage deflection angle increases from 5° to 8.32°, and then the fuselage deflection angle decreases, reaching 3.9° before the end of the ejection, and the deflection angular velocity is -1.53° / s. It can be seen that the present application has relatively accurately calculated the changes in the aircraft's heading speed, deflection angle, and deflection angular velocity over time.

[0173] Finally, it should be noted that: the drawings of the embodiments disclosed in the present invention only involve structures related to the embodiments disclosed in the present invention, and other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of the present invention can be combined with each other;

[0174] Finally: The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A method for calculating the deflection angle of a body for catapult launch when one side of the engine fails, characterized in that: include: During the catapult launch phase, a global coordinate system is established based on the aircraft's current heading, and the calculation conditions and necessary parameters for the catapult launch of the aircraft are determined by assuming that one engine fails and the other engine operates normally. In the ejection tensioning stage, all the forces on the aircraft in the current state are obtained, and all the forces are split into the resultant force in the X-axis direction, the resultant force in the Z-axis direction and the pitching moment at the center of gravity, and the static equilibrium equations of the ejection tensioning stage are established in different directions; During the ejection roll phase, obtain the rolling friction f between the aircraft's front and main landing gear tires and the ship's surface n and f m , lift L, heading resistance D, assuming that the lift and heading resistance of the aircraft act on the center of gravity of the aircraft, perform force analysis on the xoz plane in the global coordinate system, and establish the dynamic equation of the ejection run phase based on the force analysis results; In the free sliding stage, the ejection force T is cancelled, and the force analysis is performed in the X-axis direction, Z-axis direction and center of gravity position respectively, and the dynamic equation of the free sliding stage is established by combining the static equilibrium equation of the ejection tensioning stage and the dynamic equation of the ejection sliding stage; Substitute the corresponding initial conditions of each stage, solve the dynamic equations of the ejection run stage and the free run stage, and output the change of the body deflection angle data over time.

2. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure catapult launch as claimed in claim 1, characterized in that: The necessary parameters include: 1) Aircraft mass m, unit: kg; 2) Aircraft center of gravity coordinates (x) m ,y m ,z m ), unit: m; 3) The moment of inertia of the aircraft around the z-axis of the center of gravity J z , unit: kg·m 2 ; 4) Wing equivalent area S, unit: m 2 ; 5) Aircraft lift coefficient C L , dimensionless; 6) Aircraft drag coefficient C D , dimensionless; 7) Engine action point coordinates (x e ,y e ,z e ), unit: m; 8) Coordinates of the front landing gear tire contact point (x n ,y n ,z n ), unit: m; 9) Coordinates of the left main landing gear tire contact point (x ml ,y ml ,z ml ), unit: m; 10) Coordinates of the right main landing gear tire contact point (x mr ,y mr ,z mr ), unit: m; 11) Angle α between the engine thrust line and the horizontal direction, unit: °; 12) Angle β between the catapult rod and the ship deck, unit: °; 13) Angle γ between the restraining rod and the ship deck, unit: °; 14) The height of the connection point between the ejection rod and the restraining rod and the front landing gear from the ship surface, unit: m; 15) Engine thrust P, unit: N; 16) Atmospheric density ρ, unit: kg·m -3 ; 17) The rolling friction coefficient between the front landing gear tire and the ship deck is μ1, dimensionless; 18) The rolling friction coefficient between the main landing gear tire and the ship deck is μ2, dimensionless; 19) The lateral sliding friction coefficient between the main landing gear tire and the ship deck is μ3, dimensionless; 20) The relationship between the ejection force T and the stroke l provided by the catapult.

3. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: In the global coordinate system, the aircraft's forward direction is taken as the negative direction of the x-axis, the z-axis is perpendicular to the ship's surface and upward as the positive direction, and the y-axis is perpendicular to the xoz plane and points to the right side of the aircraft.

4. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: When determining the calculation conditions during the catapult takeoff phase of carrier-based aircraft, the aircraft's lateral acceleration and pitch acceleration are ignored, the impact of the aircraft carrier's swaying on the sea surface and the unevenness of the ship's surface on the aircraft load is ignored, and the impact of atmospheric disturbances on the aircraft load is ignored.

5. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: When determining the calculation conditions during the catapult takeoff phase of a carrier-based aircraft, it is assumed that the fuselage and landing gear are rigid bodies, tire compression is not considered, and the center of gravity, front landing gear, catapult rod and check rod are located on the symmetry plane of the aircraft.

6. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: During the catapult tensioning phase, the forces acting on the aircraft include: the thrust P of the right engine, the ejection force T of the ejection rod, and the tension F of the restraining rod. h , the ground reaction force N1 borne by the front landing gear, the ground reaction force N2 borne by the single main landing gear, and the gravity mg.

7. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: In the catapult tensioning stage, according to the aircraft's force in the x-axis direction being 0, the static equilibrium equation is established as: ∑F x =-Pcosα-Tcosβ-F h cosγ=0 According to the resultant force of the aircraft in the z-axis direction is 0, the static equilibrium equation is established: ∑F z =N1+2N2+Psinα-Tsinβ-F h sinγ-mg=0 According to the pitch moment of the aircraft at the center of gravity is 0, we can get: ∑M z =N1(x m -x n )-2N2(x ml -x m )+T[(-z n -h+z m )cosβ-(x m -x n )sinβ]- F h [(-z n -h+z m )cosγ-(x m -x n )sinγ]-P[(z e -z m )cosα+(x e -x m )sinα]=0。 8. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: During the catapult roll phase, the aircraft is assumed to deflect around the front main landing gear due to the unbalanced torque under the thrust of the right engine. During the deflection of the aircraft, the main landing gear tires are subjected to the sliding friction force f with the ship surface. c ;but: The equilibrium equation is established based on the resultant force in the z-axis direction: ΣF z =N1+2N2+Psinα-Tsinβ-mg+L=0 The dynamic equation is established based on the resultant force in the x-axis direction:

9. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure catapult launch as claimed in claim 2, characterized in that: The rolling friction force f n 、f m , sliding friction f c And the formula for determining the aircraft lift L and aircraft drag D is: f n =μ1N1 f m =μ2N2 <h2 style=";text-align:left;direction:ltr">f<h2 style=";text-align:left;direction:ltr"> c <h2 style=";text-align:left;direction:ltr"> =μ3N2 Where v is the aircraft heading speed.

10. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 9, characterized in that: When the aircraft deflects around the front landing gear, the dynamic equation for the aircraft fuselage deflection is: in is the total moment of the aircraft about the front wheel, is the aircraft's moment of inertia about the front wheel.

11. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure catapult launch as claimed in claim 9, characterized in that: The dynamic equation of the ejection run phase is:

12. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 2, characterized in that: The dynamic equation of the free sliding stage is: N1+2N2+Psinα-mg+L=0 J z θ(t)″=y e P-2f c (x ml -x m )。 13. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 1, characterized in that: The dynamic equations of the ejection run phase and the free run phase are numerically solved using the ode45 function in MATLAB.

14. The method for calculating the deflection angle of a launch vehicle body during a single-side engine failure ejection as claimed in claim 1, characterized in that: The body deflection angle data includes heading displacement x, heading speed x′, deflection angle θ and deflection angular speed θ′.

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