Water surface aircraft still water takeoff performance engineering estimation method

Abstract

The invention discloses a still water takeoff performance engineering estimation method for a water surface aircraft. The method comprises the following steps: (1) aerodynamic performance evaluation: calculating an aerodynamic lift coefficient, an aerodynamic resistance coefficient, a pitching moment coefficient, a transverse course static derivative, a dynamic derivative and a control derivative of the water surface aircraft; (2) hydrodynamic performance evaluation: establishing a hydrodynamic lift force calculation model, a hydrodynamic resistance / lateral force calculation model, a buoyancy estimation model and a hydrodynamic damping torque calculation model; and (3) takeoff performance evaluation: introducing a Hess driver model based on the step (1), the step (2) and engine data, and constructing a takeoff task digital virtual simulation comprehensive model in combination with a water surface aircraft still water takeoff task. The method can provide technical support for performance evaluation of static water takeoff aerodynamic force, hydrodynamic force and takeoff process of the water surface aircraft.

Description

Technical Field

[0001] The invention belongs to the technical field of theoretical mechanics, and in particular relates to an engineering estimation method for still water takeoff performance of a surface aircraft. Background Art

[0002] Water surface vehicles have the ability to take off and land on water. Compared with traditional land aircraft, they have the advantages of strong environmental adaptability and wide application. Therefore, they are widely used in both military and civilian fields. Compared with the take-off process of traditional fixed-wing aircraft on land, the high-speed gliding takeoff of water surface vehicles on the water surface is a more complex mechanical phenomenon, involving high-speed and intense actions of air, water, and solid. As the speed increases, aerodynamic force is gradually established from scratch, and hydrodynamic force is constantly changing. The pilot must maintain effective control of the aircraft's attitude and speed under the action of drastic changes in hydrodynamic force and high-intensity slipstream, which puts more stringent requirements on take-off safety. Therefore, studying the aero / hydrodynamic characteristics and take-off performance of the hydrostatic takeoff and gliding section of the water surface aircraft is a very important part of the design process of the water surface aircraft.

[0003] There is no accurate theoretical calculation method for the hydrostatic takeoff performance of surface aircraft. At present, it is mainly predicted through model tests or computational fluid dynamics methods. The prediction results of model tests are relatively accurate, but the test cost is high and has high requirements for the test environment. Although the computational fluid dynamics method can overcome the problems of the model test method, it involves the processing of large deformation of the free liquid surface and dynamic grids, and the calculation time is long and the workload is too large. Therefore, it is very necessary to propose an effective engineering estimation method for the hydrostatic takeoff performance of surface aircraft. Summary of the invention

[0004] Purpose of the Invention

[0005] In order to perform accurate theoretical calculation on the still water take-off performance of a surface aircraft, the present invention provides an engineering estimation method for the still water take-off performance of a surface aircraft.

[0006] Inventing technical solutions

[0007] The engineering estimation method for the static water takeoff performance of a surface aircraft comprises the following steps: (1) aerodynamic performance evaluation: calculating the aerodynamic lift coefficient, aerodynamic drag coefficient, pitch moment coefficient, lateral static derivative, dynamic derivative and control derivative of the surface aircraft; (2) hydrodynamic performance evaluation: establishing a hydrodynamic lift calculation model, a hydrodynamic drag / side force calculation model, a buoyancy estimation model and a hydrodynamic damping moment calculation model; (3) takeoff performance evaluation: based on step (1), step (2) and engine data, the Hess pilot model is introduced, and combined with the static water takeoff mission of the surface aircraft, a digital virtual simulation comprehensive model of the takeoff mission is constructed.

[0008] Preferably, the calculation of the aerodynamic lift coefficient includes the calculation of the airfoil lift coefficient, the wing lift coefficient, the whole aircraft lift coefficient, the flap lift coefficient, the change of the whole aircraft lift coefficient caused by the propeller slipstream, and the calculation of the lift coefficient increment caused by the ground effect, wherein the calculation method of the lift coefficient increment caused by the ground effect is: for a subsonic large aspect ratio aircraft, the angle of attack increment of the wing-body assembly is:

[0009]

[0010] Where X is the correction factor for the effect of the mirror wing tail vortex, H is the height from the ground; c r is the chord length of the wing root; b is the wing span; C Lf,WF is the lift coefficient of the wing-body assembly including the effect of flaps; C Lα,WF is the lift line slope of the wing-body combination without ground effect; r s is the finite length ground effect correction factor; is the correction factor for the effect of the attached vortex on the mirror wing; δ f is the flap deflection angle; Δ(ΔC L f) is the flap type correction factor;

[0011] Downwash increment Δε of the tail body assembly G :

[0012]

[0013] Where ε is the downwash angle without ground influence; b eff is the effective span of the wing; H H and H W are the average aerodynamic chord lengths of the horizontal tail and the wing, and the heights of the quarter-chord points from the ground;

[0014] Finally, Δα G and Δε G It is added to the angle of attack of the wing and the horizontal tail respectively, and the corresponding lift coefficient increment is the lift coefficient increment caused by the ground effect.

[0015] Preferably, the calculation of the aerodynamic drag coefficient includes the calculation of the wing drag coefficient, the fuselage drag coefficient, the whole aircraft drag coefficient, the drag coefficient of the flap, the effect of the engine on the drag, the wing induced drag increment caused by the ground effect, the calculation of the effect of the fuselage immersion, and the aerodynamic drag of the float, wherein the calculation method of the whole aircraft drag coefficient is:

[0016] Without considering the vertical tail resistance, the total aircraft drag coefficient C D It can be divided into the wing-body combination drag coefficient C DWF and the horizontal tail drag coefficient C DH Two parts:

[0017] C D =CDWF +C DH

[0018] Tail drag coefficient C DH Including the zero-lift drag coefficient C of the horizontal tail D0H and the horizontal tail lift-induced drag coefficient C DiH Two parts:

[0019]

[0020] In the formula, R HF is the horizontal tail interference factor; C fH is the friction coefficient of the flat plate surface in the horizontal tail turbulence; L' H is the maximum thickness position parameter of the horizontal tail wing, (t / c) H is the maximum relative thickness of the airfoil at the average geometric chord length of the horizontal tail; S wetH is the wetted area of ​​the horizontal tail, C LH is the horizontal tail lift coefficient; A H is the tailplane aspect ratio, e H is the spanwise factor of the horizontal tail considering viscosity;

[0021] For a subsonic high aspect ratio aircraft, the increase in wing induced drag caused by the ground effect is ΔC. Di(W,G) Determined by the following formula:

[0022]

[0023] In the formula, is the Prandtl interference factor in the multi-wing theory; is the longitudinal velocity ground correction factor; C LW is the lift coefficient of the single wing when there is no ground effect; C DW is the drag coefficient of the wing alone when there is no ground effect;

[0024] The calculation of the impact of fuselage immersion includes the calculation of fuselage friction resistance coefficient, fuselage viscous pressure resistance coefficient, and fuselage bottom resistance, among which:

[0025] The friction coefficient of the fuselage is:

[0026]

[0027] The fuselage viscous pressure drag coefficient is:

[0028]

[0029] The resistance at the bottom of the fuselage is:

[0030]

[0031] At subsonic speed, the drag caused by the fuselage lifting is small. According to the submerged area and draft volume, the linear correction algorithm is given as follows:

[0032]

[0033] Preferably, the calculation of the pitch moment coefficient includes the calculation of the wing pitch moment coefficient, the fuselage pitch moment coefficient, the whole aircraft pitch moment coefficient, and the influence of the engine on the pitch moment, wherein the influence of the water surface effect on the pitch moment is calculated as follows:

[0034] Pitching moment increment ΔC caused by water surface effect mG :

[0035] ΔC mG =ΔC mWF,G +ΔC mH,G

[0036] The pitch moment change of the wing-body assembly is related to the relative position of the center of gravity and the focus:

[0037]

[0038] The change in pitch moment of the horizontal tail:

[0039]

[0040] In the formula, is the relative position of the focus of the wing-body assembly; ΔC LWF,G is the lift increment of the wing-body assembly caused by ground effect; ΔC LH,G is the increment of the horizontal tail lift coefficient caused by ground effect; l H c is the projection of the distance from the 1 / 4 chord point of the wing to the 1 / 4 chord point of the horizontal tail on the whole aircraft coordinate system x; A is the average aerodynamic chord length; S e is the elevator area; It is the ratio of dynamic pressure at the horizontal tail to the incoming flow pressure.

[0041] Preferably, the calculation of the lateral static derivative includes the calculation of the lateral force static derivative, the rolling moment static derivative, and the yaw moment static derivative; the calculation of the dynamic derivative includes the calculation of the pitch angular velocity dynamic derivative, the angle of attack rate dynamic derivative, the roll angular velocity dynamic derivative, and the yaw angular velocity dynamic derivative; the calculation of the control derivative includes the calculation of the elevator control derivative, the aileron control derivative, and the rudder control derivative.

[0042] Preferably, the hydrodynamic lift calculation model is:

[0043]

[0044] In the formula, ρ w is the water density; r L / Bis the wetted aspect ratio; B is the fuselage width; α b =α β +k b ·(L s -L b ), α β k is the modified pitch angle considering the bottom deadrise angle β, b is the correction factor, L s is the actual wetted length of the keel, L b is the length of the straight part of the keel; r β is the wetted aspect ratio considering the bottom slope angle β.

[0045] Preferably, the hydrodynamic resistance / side force calculation model includes friction resistance, viscous pressure resistance coefficient, splash resistance, wave-making resistance, and hydrodynamic side force; wherein,

[0046] 1) Friction resistance:

[0047]

[0048] Where V1 is the average water velocity at the bottom of the fuselage, C f is the friction coefficient, ΔC f is the hull surface roughness subsidy coefficient;

[0049] 2) Viscous pressure resistance coefficient:

[0050] C v =(1+k)C f

[0051] In the formula, k is called the shape coefficient, and (1+k) is called the shape factor, which is only related to the hull shape;

[0052] 3) Splash resistance:

[0053]

[0054] Where, θ = θ / cosβ, θ is the angle between the leading edge line of the hull and the center line; α k C is the angle between the stationary point line and the center line; fs is the splash drag coefficient, taking the laminar friction drag coefficient;

[0055] 4) Wave resistance:

[0056]

[0057] Where k0 is the wave number; A(λ) is the amplitude function, f(x,z) is the half-breadth coordinate of the hull.

[0058] Preferably, the hydrodynamic damping moment calculation model includes a hydrodynamic roll damping moment, a hydrodynamic pitch damping moment, and a hydrodynamic yaw damping moment, wherein the hydrodynamic roll damping moment: L s =C lpw ρ w VλB 4 p, where C lpw is the hydrodynamic roll damping derivative; p is the roll angular velocity;

[0059] Hydrodynamic pitch (roll) damping moment: M s =C mqw ρ w Vλ 3 B 4 q, where C mqw is the hydrodynamic pitch damping derivative; q is the pitch angular velocity;

[0060] Hydrodynamic yaw (bow) damping moment: N s =C nrw ρ w VλB 4 r, where C nrw is the hydrodynamic yaw damping derivative; r is the yaw angular velocity.

[0061] Preferably, the digital virtual simulation comprehensive model of the take-off mission includes a still water take-off motion model, which includes an aircraft motion equation, an aerodynamic model, a hydrodynamic model, a mechanical control structure model, an actuator dynamics model, and an engine throttle dynamics model, wherein the aerodynamic model is established according to step (1), and the hydrodynamic model is established according to step (2); the mechanical control structure model includes a central joystick or a joystick that controls the aircraft's elevator and aileron, and the aircraft's rudder is controlled by pedals; the actuator dynamics model is an actuator deflection dynamics model of the elevator, aileron, and rudder, and the transfer function of the model is:

[0062]

[0063] In the formula, τ act is the rudder delay time constant;

[0064] The transfer function of the engine throttle dynamics model is:

[0065]

[0066] Among them, τ eng is the engine delay time constant.

[0067] Preferably, the digital virtual simulation comprehensive model of the take-off mission also includes a driver manipulation model, which includes a single-loop driver model, a multi-loop driver model, a driver perception model, a multi-channel driver model, and a still water take-off driver model; wherein the single-loop driver model is used to describe the driver's manipulation behavior under a single-channel mission, that is, it can be used to describe that the driver's manipulation target is only a single variable, and the driver does not need to consider the changes of other variables except the rate of change of the variable when controlling the variable; the multi-loop driver model is based on the single-loop model, and introduces feedback variables and derivatives of the feedback variables as outer loop feedback of the driver model; the driver model is to introduce feedback signals into the driver model;

[0068] Multi-channel driver model:

[0069] Define a multi-channel manipulation parameter dvar t , this parameter has the following relationship with the number of independent control channels n:

[0070]

[0071] Constant f and random number variance dvar v And the number of channels controlled by the driver:

[0072] f=1+10(dvar v +dvar t )

[0073] ω filter =2.4-0.4f

[0074]

[0075] In the formula, ω filter is the filter constant, which is related to the driver's ability to perceive the signal; G filter is the transfer function of the filter;

[0076] The pilot model for still water takeoff includes a pilot model for the water taxiing stage and a pilot model for the nose-up climb stage. The pilot model for the water taxiing stage contains four control channels, namely, the elevator, aileron, rudder and engine channel. The engine throttle is set to the maximum value. The elevator channel is used to control the pitch angle of the aircraft, and the aileron channel is used to ensure the roll angle; the rudder channel is mainly used to ensure the track of the aircraft when taxiing on the water; the pilot model for the nose-up climb stage contains four control channels, namely, the elevator, aileron, rudder and engine channel. The throttle is set to the maximum value. The elevator channel is used to control the pitch attitude of the aircraft to ensure that the aircraft can quickly reach a safe altitude; the rudder and aileron are jointly manipulated to control the track of the aircraft.

[0077] Advantages of the present invention: Aiming at the demand for rapid evaluation of the hydrostatic takeoff performance of surface aircraft, the present invention independently develops an engineering estimation method for hydrostatic takeoff performance of surface aircraft, which can provide technical support for the performance evaluation of hydrostatic takeoff aerodynamics, hydrodynamics and takeoff process of surface aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0078] Figure 1 It is a schematic diagram of the single-loop driver model structure.

[0079] Figure 2 It is a schematic diagram of the multi-loop driver model structure.

[0080] Figure 3 It is a schematic diagram of the driver perception model structure.

[0081] Figure 4 It is a schematic diagram of the pilot model structure during the water planing phase.

[0082] Figure 5 It is a schematic diagram of the pilot model structure during the head-up and climb phase. DETAILED DESCRIPTION

[0083] The present invention is achieved through the following technical solutions.

[0084] The engineering estimation method of the static water takeoff performance of a water surface aircraft includes the following steps:

[0085] (1) Aerodynamic performance evaluation: Based on empirical aerodynamic data and estimation formulas, the aerodynamic lift coefficient, aerodynamic drag coefficient, pitch moment coefficient, lateral static derivative, dynamic derivative, and control derivative of the surface aircraft are calculated, and an aerodynamic engineering estimation method suitable for still water takeoff of the surface aircraft is established.

[0086] 1) Aerodynamic lift coefficient

[0087] ① Airfoil lift coefficient

[0088] The lift coefficient of the airfoil is mainly determined by the zero lift angle of attack α 0,f , lift line slope C Lα,f 、The lift line deviates from the linear angle of attack α * f , maximum lift coefficient C Lmax,f and its corresponding angle of attack α max,f Since the current calculation methods cannot accurately calculate the aerodynamic characteristics after airflow separation, the parameters of the airfoil lift coefficient should be given by experiments or numerical calculations.

[0089] ②Wing lift coefficient

[0090] Maximum lift coefficient of wing C Lmax,W and its corresponding angle of attack α CLmax,W :

[0091]

[0092] In the formula, C Lmax is the maximum lift coefficient of the whole aircraft; C Lmax,f is the maximum lift coefficient of the airfoil; is the correction factor, which is determined by the sharpness of the wing leading edge; ΔC Lmax It is the Ma correction value, which is 0 at low speed; C Lα,W is the slope of the wing lift line, (A is the wing aspect ratio, β is the Prandtl–Grauward compressibility correction factor and Λ 1 / 2 is the wing 1 / 2 chord sweep angle. ); α 0,f is the zero-lift angle of attack of the airfoil; Δα CLmax It is the Ma correction value, and it is 0 at low speed.

[0093] ③ Lift coefficient of the whole aircraft

[0094]

[0095] In the formula, C Lmax,WF is the maximum lift coefficient of the wing-body fusion; C Lα,H is the slope of the horizontal tail lift line, the calculation method refers to C Lα,W ; S H is the horizontal tail area; S W is the wing area; α CLmax is the stall angle of attack; is the rate of change of the horizontal tail downwash angle with the angle of attack; ε 0,H is the tailplane downwash angle when the angle of attack is 0°; α H is the tailplane angle of attack.

[0096] ④Lift coefficient of flap

[0097] First, calculate the maximum lift coefficient increment ΔC of the trailing edge flap airfoil lmax :

[0098] ΔC lmax =K1K2K3ΔC lmax,ba

[0099] Where K1 is the flap chord length correction factor; K2 is the flap deflection correction factor; K3 is the flap motion correction factor; ΔC lmax,ba It is the basic increment of the maximum lift coefficient of the airfoil caused by the flaps.

[0100] Next, calculate the maximum coefficient increment ΔC of the wing Lmax,W :

[0101] ΔC Lmax,W=(ΔC lmax )K x (S wf / S W )

[0102] In the formula, K x is the plane shape correction factor; S wf is the flap area.

[0103] Finally, the maximum lift coefficient increment ΔC of the whole aircraft is obtained Lmax :

[0104]

[0105] In the formula, is the horizontal tail installation angle; Δε f It is the increment of the horizontal tail downwash angle caused by the flap deflection.

[0106] ⑤Effect of engine on lift

[0107] The change in lift coefficient of the whole aircraft caused by propeller slipstream is ΔC Lpow :

[0108] ΔC Lpow =ΔC LWα +ΔC LWq +ΔC LHq +ΔC LHε

[0109] Where, ΔC LWα ΔC is the increment of the wing lift coefficient caused by the downwash or upwash induced by a single propeller; LHq ΔC is the increment of wing coefficient caused by the sliding pressure of a single propeller; LHq is the lift coefficient increment caused by the dynamic pressure change at the horizontal tail; ΔC LHε is the lift coefficient increment caused by the downwash change of the horizontal tail.

[0110] ⑥Water surface effect

[0111] For subsonic high aspect ratio aircraft, the angle of attack increment of the wing-body assembly is:

[0112]

[0113] Where X is the correction factor for the effect of the mirror wing tail vortex, which is given by From the table, we can see that H is the height from the ground; c r is the chord length of the wing root; b is the wing span; C Lf,WF is the lift coefficient of the wing-body assembly including the effect of flaps; C Lα,WF is the lift line slope of the wing-body combination without ground effect; r s is the finite length ground effect correction factor; is the correction factor for the effect of the attached vortex on the mirror wing; δ f is the flap deflection angle; Δ(ΔC L f) is the flap form correction factor.

[0114] Downwash increment Δε of the tail body assembly G :

[0115]

[0116] Where ε is the downwash angle without ground influence; b eff is the effective span of the wing; H H and H W They are the heights from the ground to the 1 / 4 chord point of the average aerodynamic chord length of the horizontal tail and the wing respectively.

[0117] Finally, Δα G and Δε G It is added to the angle of attack of the wing and the horizontal tail respectively, and the corresponding lift coefficient increment is the lift coefficient increment caused by the ground effect.

[0118] 2) Aerodynamic drag coefficient

[0119] ① Wing drag coefficient

[0120] Wing drag coefficient C DW It can be divided into the wing zero lift drag coefficient C D0W and the wing lift-drag coefficient C DiW Two parts:

[0121] C DW =C D0W +C DiW

[0122]

[0123] In the formula, R WF is the wing interference factor, for a single wing R WF =1; R LS is the lifting surface correction factor, which is given by Ma number and sweep angle according to the empirical method; C fW is the turbulent friction coefficient of the wing, which is calculated from the Reynolds number Re and the Ma number; L′ is the maximum thickness position parameter of the airfoil, which is 1.2 when the maximum thickness position of the airfoil is greater than 0.3c, otherwise it is 2; t / c is the maximum relative thickness of the airfoil at the average geometric chord length of the wing; S wetW is the wetted area of ​​the wing. LW is the lift coefficient of the wing; e is the spanwise factor considering viscosity; τ W is the linear torsion angle, and the inner wash is positive; v is the induced drag factor of linear torsion; ω is the zero-lift drag factor of linear torsion.

[0124] ② Fuselage drag coefficient

[0125] Fuselage drag coefficient C DF It can be divided into the zero-lift drag coefficient C D0F and fuselage lift-induced drag coefficient C DiF Two parts:

[0126] C DF =C D0F +C DiF

[0127] Zero lift drag coefficient C D0F Mainly includes friction resistance coefficient C DfF , Viscous pressure resistance coefficient C DpF and the bottom resistance coefficient C DbF :

[0128] C D0F =C DfF +C DpF +C DbF

[0129] C DfF =R WF C fF S wetF / S W

[0130] C DpF =R WF C fF [60 / (l F / d F ) 3 +0.0025(l F / d F )]S wetF / S W

[0131] C DbF ={0.029(d b / d F ) 3 / [(C DfF +C DpF )(S W / S F ) 1 / 2 ]}(S F / S W )

[0132] In the formula, R WF is the fuselage interference factor, which is a function of the Ma number and the Re number, and is 1 for a single fuselage; C fF is the friction coefficient of the turbulent flat surface of the fuselage, which is determined according to the different surface materials of the fuselage and the Ma number and Re number; SwetF is the wetted area of ​​the fuselage; l F is the reference length of the fuselage; d F is the maximum diameter of the fuselage; d b is the bottom diameter of the fuselage; S F is the fuselage area.

[0133] Fuselage lift-induced drag coefficient C DiF Determined by the following formula:

[0134]

[0135] Where α is the aircraft angle of attack; V F is the volume of the fuselage; η is the ratio of the resistance of the finite cylinder to the resistance of the infinite cylinder, and the engineering experience value is taken according to the geometric dimensions of the fuselage; C Dc is the steady-state cross-flow resistance coefficient of the original column of infinite length.

[0136] ③ Overall drag coefficient

[0137] Without considering the vertical tail resistance, the total aircraft drag coefficient C D It can be divided into the wing-body combination drag coefficient C DWF and the horizontal tail drag coefficient C DH Two parts:

[0138] C D =C DWF +C DH

[0139] The drag calculation of the wing-body assembly is to give the experimental interference correction factor based on the calculated wing and fuselage drag, and does not consider the interference of the fuselage on the horizontal tail drag.

[0140] Tail drag coefficient C DH Including the zero-lift drag coefficient C of the horizontal tail D0H and the horizontal tail lift-induced drag coefficient C DiH Two parts:

[0141]

[0142] In the formula, R HF is the horizontal tail interference factor; C fH is the friction coefficient of the flat plate surface in the horizontal tail turbulence; L' H is the maximum thickness position parameter of the horizontal tail wing. (t / c) H is the maximum relative thickness of the airfoil at the average geometric chord length of the horizontal tail; S wetH is the wetted area of ​​the horizontal tail. LH is the horizontal tail lift coefficient; A H is the tailplane aspect ratio, e H is the spanwise factor of the horizontal tail considering viscosity.

[0143] ④ Drag coefficient of flap

[0144] When the flaps are lowered, the additional drag generated is composed of three parts: the incremental drag ΔC Dprof , lift-induced resistance increment ΔC Di and the disturbance resistance increment ΔC Dint .

[0145] ΔC Dflap =ΔC Dprof +ΔC Di +ΔC Dint

[0146] ΔC Dprof =(ΔC dp,Λ1 / 4＝0 )cosΛ 1 / 4 S wf / S W

[0147] ΔC Di =K 2 (ΔC Lflap ) 2 cosΛ 1 / 4

[0148] ΔC Dint =K int ΔC Dprof

[0149] Where, ΔC dp,Λ1 / 4＝0 is the drag increment of the two-dimensional lift-enhancing device; 1 / 4 ΔC is the 1 / 4 chord sweep angle. Lflap is the lift increment caused by the high-lift device; K int is the type factor of the high-lift device, the simple flap is 0 and the single-slotted flap is 0.4.

[0150] ⑤Effect of engine on resistance

[0151] The effect of the engine on drag includes the zero-lift drag increment ΔC Dopow and lift-induced drag coefficient increment ΔC Dipow .

[0152] ΔC Dpow =ΔC D0pow +ΔC Dipow

[0153] ⑥The influence of water surface effect on resistance

[0154] For a subsonic high aspect ratio aircraft, the increase in wing induced drag caused by the ground effect is ΔC. Di(W,G) It can be determined by the following formula:

[0155]

[0156] In the formula, is the Prandtl interference factor in the multi-wing theory; is the longitudinal velocity ground correction factor; C LW is the lift coefficient of the single wing when there is no ground effect; C DW is the drag coefficient of the wing alone when out of ground effect;

[0157] ⑦ Impact of water immersion on the fuselage

[0158] The friction resistance of the fuselage mainly depends on its surface roughness and surface area. When the fuselage is immersed in water, the surface area exposed to the air decreases, and the friction resistance coefficient of the fuselage is:

[0159]

[0160] The fuselage viscous pressure drag coefficient is:

[0161]

[0162] Fuselage bottom drag is the drag caused by the airflow separating at the bottom of the fuselage, and is therefore related to the bottom area:

[0163]

[0164] At subsonic speed, the drag caused by the fuselage lifting is small, and the linear correction algorithm can be given according to the submerged area and draft volume as follows:

[0165]

[0166] ⑧ Aerodynamic resistance of the buoy

[0167] After the geometric parameters of the buoy are modeled, the model is constructed based on a semi-theoretical and semi-empirical formula.

[0168] 3) Pitch moment coefficient

[0169] ① Wing pitch moment coefficient

[0170] The point of action of the lift increment generated by the change in the angle of attack of the wing is called the aerodynamic center (focus). The pitch moment of the wing about the focus does not change with the angle of attack. This pitch moment is called the zero-lift pitch moment C of the wing. m0,W Part of it comes from the airfoil, and the other part comes from the torsion and plane shape of the wing. When the entire wing is at zero lift angle of attack, the lift of the local section is not all zero, which will cause additional zero lift moment. For a large aspect ratio wing, the zero lift moment can be calculated by the following formula:

[0171]

[0172] In the formula, C m0,r and C m0,tare the zero-lift pitching moment coefficients of the wing root section and the wing tip section respectively; is the increment of zero lift pitching moment produced per unit twist angle.

[0173] ②Fuselage pitch moment coefficient

[0174] The fuselage pitch moment coefficient is composed of two parts: the front flow contribution of the fuselage and the rear viscous cross flow contribution:

[0175]

[0176] Where K2-K1 is the apparent quality factor, which is the fuselage length l F and fuselage equivalent diameter d F The function of the ratio of F is the effective angle of attack of the fuselage, α F =α-α 0F , α 0F A is the zero lift angle of attack of the fuselage; F is the maximum cross-sectional area of ​​the fuselage; x0 is the coordinate of the end point of the fuselage potential flow, which is related to the shape and size of the fuselage; A x is the cross-sectional area at any position x of the fuselage; r x is the radius at any position x of the fuselage; ref is the coordinate of the torque reference point.

[0177] ③Full aircraft pitch moment coefficient

[0178] Normal layout aircraft only considers the contribution of the horizontal tail to the pitching moment, and the horizontal tail pitching moment coefficient C considering the influence of the fuselage mHF :

[0179]

[0180] Taking into account the effects of speed retardation and downwash, the horizontal tail lift coefficient C LHF It can be expressed as:

[0181]

[0182] In the formula, and is the relative position of the aircraft's center of gravity; and is the focus position of the whole machine; C DHF is the horizontal tail drag coefficient; k q is the dynamic pressure ratio at the horizontal tail; K G K is the influence factor of the center of gravity position on downwash; αH K is the ratio coefficient between the wing-body lift and the single exposed wing lift when the horizontal tail and the fuselage are at the same angle of attack; δ0His the ratio coefficient between the interfering wing-body lift and the single exposed wing lift when the fuselage angle of attack is zero and the horizontal tail has an installation angle; Mounting angle for the horizontal tail.

[0183] ④Effect of engine on pitch moment

[0184] Engine contribution to total pitch moment ΔC m :

[0185] ΔC m =ΔC m0,T +ΔC m0,N +ΔC m,pow

[0186] Where, ΔC m0,T ΔC is the zero-lift pitch moment increment when the engine thrust line does not pass the center of gravity; m0,N is the zero-lift pitch moment increment caused by the propeller normal force; ΔC m,pow is the total effect of slipstream on the pitch moment of the entire aircraft.

[0187] ⑤The influence of water surface effect on pitching moment

[0188] Pitching moment increment ΔC caused by water surface effect mG :

[0189] ΔC mG =ΔC mWF,G +ΔC mH,G

[0190] The pitch moment change of the wing-body assembly is related to the relative position of the center of gravity and the focus:

[0191]

[0192] The change in pitch moment of the horizontal tail:

[0193]

[0194] In the formula, is the relative position of the focus of the wing-body assembly; ΔC LWF,G is the lift increment of the wing-body assembly caused by ground effect; ΔC LH,G is the increment of the horizontal tail lift coefficient caused by ground effect; l H c is the projection of the distance from the 1 / 4 chord point of the wing to the 1 / 4 chord point of the horizontal tail on the whole aircraft coordinate system x; A is the average aerodynamic chord length; S e is the elevator area; It is the ratio of dynamic pressure at the horizontal tail to the incoming flow pressure.

[0195] 4) Lateral static derivative

[0196] ① Static derivative of lateral force

[0197] The derivative of the lateral force coefficient of the whole aircraft with respect to the sideslip angle C Yβ :

[0198] C Yβ =C Yβ,WF +C YβV

[0199] The derivative of the side force coefficient of the wing-body assembly with respect to the sideslip angle C Yβ,WF :

[0200]

[0201] Vertical tail pair C Yβ The contribution includes the interference effect of the fuselage on the vertical tail:

[0202]

[0203] In the formula, is the ratio of the maximum cross-sectional area of ​​the fuselage to the wing area, β is the sideslip angle; is the ratio of the fuselage side area to the wing area; K V is an empirical factor, given according to the shape and size of the vertical tail; C LαV is the slope of the vertical tail lift line; σ is the vertical tail side wash; S V is the vertical tail area; Z W is the vertical height of the fuselage at the root chord of the wing; Z F is the maximum fuselage depth.

[0204] ②Static derivative of rolling moment

[0205] The derivative of the rolling moment coefficient of the whole aircraft with respect to the sideslip angle C lβ :

[0206] C lβ =C lβWF +C lβV

[0207] Contribution of the wing-body assembly to the rolling moment derivative C lβWF :

[0208]

[0209] Vertical tail pair C lβ Contribution of C lβV :

[0210]

[0211] In the formula, C LWF is the lift coefficient of the wing-body assembly; is the wing sweep contribution, obtained by interpolation based on experimental and empirical values; K MΛ is the compressibility correction for the sweep; K F is the fuselage correction factor; is the aspect ratio contribution, obtained by interpolation based on experimental and empirical values; Γ is the geometric dihedral angle of the wing; is the wing dihedral effect, obtained by interpolation based on experimental and empirical values; K MΓ Compressibility correction for the upthrust contribution; is the correction of the inductive effect of the fuselage on the wing height; ΔC lβ,ZW is the increment of the rolling moment derivative at different positions of the wing relative to the fuselage; is the wing twist correction factor; and ΔC lβ,ZW Can be obtained by Find, d F av = (average cross-sectional area of ​​fuselage / 0.7854) 1 / 2 ; l V Z is the horizontal distance from the center of gravity of the aircraft to the aerodynamic center of the vertical tail; V It is the vertical distance from the center of gravity of the aircraft to the aerodynamic center of the vertical tail.

[0212] ③Static derivative of yaw moment

[0213] C nβ =C nβWF +C nβV

[0214] The yaw moment derivative of the wing-body assembly in the subsonic small angle of attack range:

[0215]

[0216] Contribution of the vertical tail to the yaw moment derivative:

[0217]

[0218] In the formula, K WF is the wing-body interference factor, obtained by interpolation based on experimental and empirical results; K RF is the correction factor for the Reynolds number, obtained by interpolation based on experimental and empirical results; S FS is the side area of ​​the fuselage; l F is the fuselage length.

[0219] 5) Dynamic derivative

[0220] ① Pitch angular velocity dynamic derivative

[0221] The dynamic derivative of the lift coefficient caused by the pitch angular velocity of the whole aircraft C Lq and the dynamic derivative of the pitch moment coefficient C mq :

[0222]

[0223] In the formula, is the dimensionless pitch angular velocity.

[0224] ② Angle of attack rate derivative

[0225] Angle of attack rate The dynamic derivatives caused include the dynamic derivative of the lift coefficient and the dynamic derivative of the pitch moment coefficient

[0226]

[0227] In the formula, is the dimensionless rate of change of angle of attack.

[0228] ③ Dynamic derivative of rolling angular velocity

[0229] The dynamic derivative of the rolling angular velocity includes the dynamic derivative of the rolling moment coefficient C lp and the dynamic derivative of the yaw moment coefficient C np :

[0230]

[0231] In the formula, For a certain number of Ma, C L = 0, the combined parameters of the dynamic derivative of the wing rolling moment coefficient are given by looking up the table based on engineering experience values; C lpW,Γ C is caused by the wing upward lp Component; ΔC lp,drag C is the drag caused by the wing lp Quantity; C is the lift caused by the wing np contribute; C is caused by the torsion angle np contribute; C is caused by flap deflection np Contribution; C npΓ C is caused by the wing sweep angle np contribute.

[0232] ④ Dynamic derivative of yaw angular velocity

[0233] The lateral dynamic derivatives caused by the yaw angular velocity include C lr and C nr :

[0234]

[0235] The terms in the formula are the wing planform, dihedral angle, twist angle, flap deflection and vertical tail C. lrcontribution.

[0236]

[0237] 6) Manipulating derivatives

[0238] ① Elevator control derivative

[0239] The elevator control derivatives include the lift coefficient and the pitch moment coefficient with respect to the elevator deflection angle C Lδe and C mδe .

[0240] C Lδe =α δe C LδH

[0241] C mδe =α δe C mδH

[0242] In the formula, α δe is the rudder efficiency; C LδH is the derivative of the lift coefficient with respect to the horizontal tail; C mδH is the derivative of the lift moment coefficient with respect to the horizontal tail.

[0243] ② Aileron control derivative

[0244] Aileron deflection will cause changes in the aircraft's rolling moment and yaw moment, and the corresponding control derivative is C lδa and C nδa .

[0245] C lδa =C lδa,le +C lδa,ri

[0246] In the formula, C lδa,le and C lδa,ri Represent the roll control derivatives of the left and right ailerons respectively.

[0247] C nδa =K a C LW C lδa

[0248] In the formula, K a is the correction constant, which is obtained by looking up the table according to the aileron geometric parameters; C LW is the wing lift coefficient.

[0249] ③Rudder control derivative

[0250] The rudder control derivatives include the lateral force coefficient derivative C Yδr , Rolling moment coefficient derivative C lδr and the derivative of the yaw moment coefficient C nδr .

[0251]

[0252] C lδr =C Yδr (Z Vcosα -l V sinα) / b

[0253] C nδr =-C Yδr (l V cosα+Z V sinα) / b

[0254] (2) Hydrodynamic performance evaluation: Based on the lift theory of planing boats, the flat plate lift calculation formula is modified to establish a hydrodynamic lift calculation model; refer to the empirical formula for ship resistance to establish a hydrodynamic resistance / side force calculation model. A buoyancy estimation model is established based on Archimedes' principle; a hydrodynamic damping torque calculation model is established based on the empirical calculation formula for ship damping torque; and a hydrodynamic engineering estimation method suitable for still water takeoff of surface aircraft is established.

[0255] 1) Hydrodynamic lift

[0256]

[0257] In the formula, ρ w is the water density; r L / B is the wetted aspect ratio; B is the fuselage width; α b =α β +k b ·(L s -L b ), α β k is the modified pitch angle considering the bottom deadrise angle β, b is the correction factor, L s is the actual wetted length of the keel, L b is the length of the straight part of the keel; r β is the wetted aspect ratio considering the bottom slope angle β.

[0258] 2) Hydrodynamic resistance

[0259] ① Friction resistance

[0260]

[0261] Where V1 is the average water velocity at the bottom of the fuselage, C f is the friction coefficient, ΔC f is the hull surface roughness compensation coefficient, take ΔC f =0.4×10 -3 .

[0262] ②Viscous pressure resistance coefficient

[0263] C v =(1+k)C f

[0264] In the formula, k is called the shape coefficient and (1+k) is called the shape factor, which is only related to the hull shape.

[0265] ③Splash resistance

[0266]

[0267] Where, θ = θ / cosβ, θ is the angle between the leading edge line of the hull and the center line; α k C is the angle between the stationary point line and the center line; fs is the splash drag coefficient, and is the laminar friction drag coefficient.

[0268] ④ Wave-making resistance

[0269]

[0270] Where k0 is the wave number; A(λ) is the amplitude function, f(x,z) is the half-breadth coordinate of the hull.

[0271] 3) Hydrodynamic lateral force

[0272]

[0273] Where, u is the component of the aircraft ground speed on the X-axis of the aircraft coordinate system; β f is the sideslip angle of the aircraft.

[0274] 4) Buoyancy

[0275] According to Archimedes' principle:

[0276] F bu =ρ w V p

[0277] Where V p is the displacement volume.

[0278] 5) Hydrodynamic damping torque

[0279] ①Hydrodynamic roll (roll) damping torque

[0280] L s =C lpw ρ w VλB 4 p

[0281] In the formula, C lpwis the hydrodynamic roll damping derivative; p is the roll angular velocity.

[0282] ②Hydrodynamic pitch (roll) damping moment

[0283] M s =C mqw ρ w Vλ 3 B 4 q

[0284] In the formula, C mqw is the hydrodynamic pitch damping derivative; q is the pitch angular velocity.

[0285] ③Hydrodynamic yaw (bow roll) damping moment

[0286] N s =C nrw ρ w VλB 4 r

[0287] In the formula, C nrw is the hydrodynamic yaw damping derivative; r is the yaw angular velocity.

[0288] (3) Take-off performance evaluation: Based on the established aerodynamic / hydrodynamic engineering estimation method and engine data, the Hess pilot model is introduced, and combined with the surface aircraft still water take-off mission, a comprehensive digital virtual simulation model of the take-off mission is constructed.

[0289] 1) Still water takeoff motion model

[0290] ① Aircraft motion equation

[0291] In an inertial reference frame, the dynamic equation for the translation of the center of mass of a rigid body is:

[0292]

[0293] In the formula, m b F is the mass of the aircraft body; T is the pulling force of a single engine; n T is the number of working engines; F is the engine pitch installation angle; w It is the force provided by the water surface when the aircraft is gliding on the water surface; D, C, L are the aerodynamic drag, side force and lift components defined along the airflow reference system.

[0294] In the body reference system, the rotational dynamics equation of a rigid body is:

[0295]

[0296] Where p, q, and r are the three-axis angular velocities of the body axis system; I x ,Iy ,I z are the moments of inertia of the three axes of the aircraft; I zx is the inertia product of the aircraft; y Ti ,z Ti is the coordinates of the thrust action point of the i-th engine; M W is the torque provided by the water surface; L, M, N are the components of the aerodynamic torque on the body axis system.

[0297] ② Aerodynamic model and hydrodynamic model

[0298] For the aerodynamic model, see “(1) Aerodynamic performance evaluation”; for the hydrodynamic model, see “(2) Hydrodynamic performance evaluation”.

[0299] ③Mechanical control structure model

[0300] The main control system of the aircraft includes the control wheel / stick and pedals. In the mechanical control system, the control device generally uses a central control stick or control wheel to control the aircraft's elevator and ailerons, while the pedals control the aircraft's rudder.

[0301] ④Actuator dynamics model

[0302] The dynamic characteristics of the actuator are mainly determined by the hydraulic booster. When modeling and simulating, the actuator deflection dynamics model of the elevator, aileron and rudder can be described by the first-order inertial link, and its transfer function is:

[0303]

[0304] In the formula, τ act is the delay time constant of the control surface. In the simulation modeling, considering the position constraint and deflection rate limit of the control surface, a saturation limit link is added.

[0305] ⑤Engine model

[0306] The engine throttle dynamics model is a first-order inertia link, and its transfer function is:

[0307]

[0308] Among them, τ eng is the engine delay time constant. When establishing the engine throttle dynamics model, the saturation characteristic link is adopted considering the power limitation of the engine.

[0309] 2) Driver operation model

[0310] ①Single-loop driver model

[0311] The single-loop driver model is used to describe the driver's manipulation behavior under a single-channel task, that is, it can be used to describe that the driver's manipulation target is only a single variable, and the driver does not need to consider the changes of other variables except the rate of change of the variable when controlling the variable. Figure 1 , where M represents the flight motion variable that the pilot expects to control; is the rate of change of the desired control variable; C is the driver's control instruction to M, which is generally determined by the task; G nm Represents a highly simplified model of the driver's muscle dynamics (leg or arm).

[0312] ②Multi-circuit driver model

[0313] The multi-loop driver model can be regarded as an extension of the single-loop, that is, a new loop is introduced on the basis of the single-loop driver model to construct a more complex driver model. In general, actual control tasks often require the driver to control more state variables. Taking tracking altitude as an example, the driver often needs to pay attention to the control of pitch attitude while tracking altitude. The multi-loop driver model is used to describe the driver's control under such tasks. The multi-loop model is based on the single-loop model and introduces feedback variables and their derivatives as the outer loop feedback of the driver model. The multi-loop driver model is as follows: Figure 2 .

[0314] ③Driver perception model

[0315] like Figure 3 ,In the pilot model, the pilot model needs to perceive the aircraft's state information, that is, ,feedback signals are introduced into the pilot model.

[0316] ④Multi-channel driver model

[0317] Since the aircraft propeller engine will cause a certain amount of air wash to the wing, and the flight mission may be affected by factors such as wind and waves, multiple channels should be taken into account during water takeoff and landing. Define a multi-channel control parameter dvar t , this parameter has the following relationship with the number of independent control channels n:

[0318]

[0319] Constant f and random number variance dvar v And the number of channels controlled by the driver:

[0320] f=1+10(dvar v +dvar t )

[0321] ω filter=2.4-0.4f

[0322]

[0323] In the formula, ω filter is the filter constant, which is related to the driver's ability to perceive the signal; G filter is the transfer function of the filter.

[0324] ⑤Still water takeoff pilot model

[0325] The structure of the pilot model during the water planing phase is as follows Figure 4 As shown in the figure, it contains four control channels, namely, elevator, aileron, rudder and engine channels, and the engine throttle is set to the maximum value. The elevator channel is mainly used to control the pitch angle of the aircraft, that is, the longitudinal inclination angle of the aircraft when gliding on the water surface. The aileron channel is mainly used to ensure that the roll angle is not too large. The rudder channel is mainly used to ensure that the track of the aircraft will not deviate greatly when gliding on the water surface.

[0326] The pilot model structure in the head-up climbing phase is as follows Figure 5 As shown in the figure, it contains four control channels, namely elevator, aileron, rudder and engine channels, and the throttle is set to the maximum value. The elevator channel is mainly used to control the pitch attitude of the aircraft to ensure that the aircraft can quickly reach a safe altitude. The rudder and aileron are combined to control the aircraft's track to ensure that the aircraft will not have a large track deviation.

[0327] The protection scope of the present invention is not limited to the above-mentioned embodiments. Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the scope of the present invention. If these changes and modifications fall within the scope of the claims of the present invention and their equivalents, the present invention is intended to include these changes and modifications.

Claims

1. A method for estimating the performance of a surface aircraft taking off from water, characterized in that: The steps include: (1) Aerodynamic performance evaluation: calculation of the aerodynamic lift coefficient, aerodynamic drag coefficient, pitch moment coefficient, lateral static derivative, dynamic derivative and control derivative of the water surface aircraft; (2) Hydrodynamic performance evaluation: Establish hydrodynamic lift calculation model, hydrodynamic resistance / side force calculation model, buoyancy estimation model, and hydrodynamic damping moment calculation model; (3) Take-off performance evaluation: Based on step (1), step (2) and engine data, the Hess pilot model is introduced, and combined with the surface aircraft still water take-off mission, a comprehensive digital virtual simulation model of the take-off mission is constructed.

2. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 1, characterized in that: The calculation of aerodynamic lift coefficient includes the calculation of airfoil lift coefficient, wing lift coefficient, whole aircraft lift coefficient, flap lift coefficient, change of whole aircraft lift coefficient caused by propeller slipstream, and lift coefficient increment caused by ground effect. The calculation method of lift coefficient increment caused by ground effect is: For subsonic high aspect ratio aircraft, the angle of attack increment of the wing-body assembly is: Where X is the correction factor for the effect of the mirror wing tail vortex, H is the height from the ground; c r is the chord length of the wing root; b is the wing span; C Lf,WF is the lift coefficient of the wing-body assembly including the effect of flaps; C Lα,WF is the lift line slope of the wing-body combination without ground effect; r s is the finite length ground effect correction factor; is the correction factor for the effect of the attached vortex on the mirror wing; δ f is the flap deflection angle; Δ(ΔC L f) is the flap type correction factor; Downwash increment Δε of the tail body assembly G : Where ε is the downwash angle without ground influence; b eff is the effective span of the wing; H H and H W are the average aerodynamic chord lengths of the horizontal tail and the wing, and the heights of the quarter-chord points from the ground; Finally, Δα G and Δε G It is added to the angle of attack of the wing and the horizontal tail respectively, and the corresponding lift coefficient increment is the lift coefficient increment caused by the ground effect.

3. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 1, characterized in that: The calculation of aerodynamic drag coefficient includes wing drag coefficient, fuselage drag coefficient, whole aircraft drag coefficient, flap drag coefficient, engine effect on drag, wing induced drag increment caused by ground effect, calculation of fuselage immersion effect, and aerodynamic drag of floats. The calculation method of whole aircraft drag coefficient is: Without considering the vertical tail resistance, the total aircraft drag coefficient C D It can be divided into the wing-body combination drag coefficient C DWF and the horizontal tail drag coefficient C DH Two parts: C D =C DWF +C DH Tail drag coefficient C DH Including the zero-lift drag coefficient C of the horizontal tail D0H and the horizontal tail lift-induced drag coefficient C DiH Two parts: In the formula, R HF is the horizontal tail interference factor; C fH is the friction coefficient of the flat plate surface in the horizontal tail turbulence; L' H is the maximum thickness position parameter of the horizontal tail wing, (t / c) H is the maximum relative thickness of the airfoil at the average geometric chord length of the horizontal tail; S wetH is the wetted area of ​​the horizontal tail, C LH is the horizontal tail lift coefficient; A H is the tailplane aspect ratio, e H is the spanwise factor of the horizontal tail considering viscosity; For a subsonic high aspect ratio aircraft, the increase in wing induced drag caused by the ground effect is ΔC. Di(W,G) Determined by the following formula: In the formula, is the Prandtl interference factor in the multi-wing theory; is the longitudinal velocity ground correction factor; C LW is the lift coefficient of the single wing when there is no ground effect; C DW is the drag coefficient of the wing alone when there is no ground effect; The calculation of the impact of fuselage immersion includes the calculation of fuselage friction resistance coefficient, fuselage viscous pressure resistance coefficient, and fuselage bottom resistance, among which: The friction coefficient of the fuselage is: The fuselage viscous pressure drag coefficient is: The bottom resistance of the fuselage is: At subsonic speed, the drag caused by the fuselage lifting is small. According to the submerged area and draft volume, the linear correction algorithm is given as follows:

4. The method for engineering estimation of still water takeoff performance of a watercraft according to claim 1, characterized in that: The calculation of the pitch moment coefficient includes the calculation of the wing pitch moment coefficient, the fuselage pitch moment coefficient, the whole aircraft pitch moment coefficient, and the effect of the engine on the pitch moment. Among them, the effect of the water surface effect on the pitch moment is calculated as follows: Pitching moment increment ΔC caused by water surface effect mG : ΔC mG =ΔC mWF,G +ΔC mH,G The pitch moment change of the wing-body assembly is related to the relative position of the center of gravity and the focus: The change in pitch moment of the horizontal tail: In the formula, is the relative position of the focus of the wing-body assembly; ΔC LWF,G is the lift increment of the wing-body assembly caused by ground effect; ΔC LH,G is the increment of the horizontal tail lift coefficient caused by ground effect; l H c is the projection of the distance from the 1 / 4 chord point of the wing to the 1 / 4 chord point of the horizontal tail on the whole aircraft coordinate system x; A is the average aerodynamic chord length; S e is the elevator area; It is the ratio of dynamic pressure at the horizontal tail to the incoming flow pressure.

5. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 1, characterized in that: The calculation of lateral static derivatives includes the calculation of lateral force static derivative, rolling moment static derivative, and yaw moment static derivative; the calculation of dynamic derivatives includes the calculation of pitch angular velocity dynamic derivative, angle of attack rate dynamic derivative, roll angular velocity dynamic derivative, and yaw angular velocity dynamic derivative; the calculation of control derivatives includes the calculation of elevator control derivative, aileron control derivative, and rudder control derivative.

6. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 1, characterized in that: The hydrodynamic lift calculation model is: In the formula, ρ w is the water density; r L / B is the wetted aspect ratio; B is the fuselage width; α b =α β +k b ·(L s -L b ), α β k is the modified pitch angle considering the bottom deadrise angle β, b is the correction factor, L s is the actual wetted length of the keel, L b is the length of the straight part of the keel; r β is the wetted aspect ratio considering the bottom slope angle β.

7. The method for estimating the performance of a watercraft taking off from still water as claimed in claim 1, characterized in that: The hydrodynamic resistance / side force calculation model includes friction resistance, viscous pressure resistance coefficient, splash resistance, wave resistance, and hydrodynamic side force; among them, 1) Friction resistance: R f =0.5ρ w V1 2 S wetF (C f +ΔC f ) Where V1 is the average water velocity at the bottom of the fuselage, C f is the friction coefficient, ΔC f is the hull surface roughness subsidy coefficient; 2) Viscous pressure resistance coefficient: C v =(1+k)C f In the formula, k is called the shape coefficient, and (1+k) is called the shape factor, which is only related to the hull shape; 3) Splash resistance: Where, θ = θ / cosβ, θ is the angle between the leading edge line of the hull and the center line; α k C is the angle between the stationary point line and the center line; fs is the splash drag coefficient, taking the laminar friction drag coefficient; 4) Wave resistance: Where k0 is the wave number; A(λ) is the amplitude function, f(x,z) is the half-breadth coordinate of the hull.

8. The method for estimating the performance of a watercraft taking off from still water as claimed in claim 1, characterized in that: The calculation model of the hydrodynamic damping moment includes the hydrodynamic roll damping moment, the hydrodynamic pitch damping moment, and the hydrodynamic yaw damping moment, among which: Hydrodynamic roll damping torque: L s =C lpw r w VλB 4 p In the formula, C lpw is the hydrodynamic roll damping derivative; p is the roll angular velocity; Hydrodynamic pitch (roll) damping moment: M s =C mqw r w Vλ 3 B 4 q In the formula, C mqw is the hydrodynamic pitch damping derivative; q is the pitch angular velocity; Hydrodynamic yaw (bow) damping torque: N s =C nrw r w VλB 4 r In the formula, C nrw is the hydrodynamic yaw damping derivative; r is the yaw angular velocity.

9. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 1, characterized in that: The digital virtual simulation comprehensive model of the take-off mission includes a static water take-off motion model, which includes an aircraft motion equation, an aerodynamic model, a hydrodynamic model, a mechanical control structure model, an actuator dynamic model, and an engine throttle dynamic model, wherein the aerodynamic model is established according to step (1), and the hydrodynamic model is established according to step (2); The mechanical control structure model includes a central control stick or control wheel to control the aircraft's elevator and ailerons, while the aircraft's rudder is controlled by pedals; The actuator dynamics model is the actuator deflection dynamics model of the elevator, aileron and rudder, and the transfer function of the model is: In the formula, τ act is the delay time constant of the rudder surface; The transfer function of the engine throttle dynamics model is: Among them, τ eng is the engine delay time constant.

10. The method for estimating the performance of a water surface vehicle for hydrostatic takeoff according to claim 9, characterized in that: The digital virtual simulation comprehensive model of the take-off mission also includes a pilot control model, which includes a single-loop pilot model, a multi-loop pilot model, a pilot perception model, a multi-channel pilot model, and a still water take-off pilot model. Among them, the single-loop driver model is used to describe the driver's manipulation behavior under a single-channel task, that is, it can be used to describe that the driver's manipulation target is only a single variable, and the driver does not need to consider the changes of other variables except the change rate of the variable when controlling the variable; The multi-loop driver model is based on the single-loop model, and introduces feedback variables and the derivatives of feedback variables as the outer loop feedback of the driver model; The driver model is to introduce feedback signals into the driver model; Multi-channel driver model: Define a multi-channel manipulation parameter dvar t , this parameter has the following relationship with the number of independent control channels n: Constant f and random number variance dvar v And the number of channels controlled by the driver: f=1+10(dvar v +dvar t , oh filter =2.4-0.4f In the formula, ω filter is the filter constant, which is related to the driver's ability to perceive the signal; G filter is the transfer function of the filter; The pilot model for still water takeoff includes the pilot model for the taxiing stage and the pilot model for the head-up climb stage. The pilot model for the taxiing stage contains four control channels, namely the elevator, aileron, rudder and engine channels. The engine throttle is set to the maximum value. The elevator channel is used to control the pitch angle of the aircraft, and the aileron channel is used to ensure the roll angle; the rudder channel is mainly used to ensure the track of the aircraft when taxiing on the water. The pilot model in the head-up climb phase contains four control channels, namely the elevator, aileron, rudder and engine channels. The throttle is set to the maximum value. The elevator channel is used to control the pitch attitude of the aircraft to ensure that the aircraft can reach a safe altitude quickly; the rudder and aileron are jointly controlled to control the aircraft's track.

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