Conveyor thrust-weight ratio design method and device

Abstract

The invention belongs to the technical field of flight control, and particularly relates to a transport plane thrust-weight ratio design method and device which are applied to a lower surface blowing flap power high lift transport plane. The method comprises the steps that S1, transport wing loads under different constraint conditions are calculated, and the minimum value of the transport wing loads is taken as the design wing load of the lower surface blowing flap dynamic high-lift transport plane; s2, calculating a first thrust-weight ratio with a takeoff and taxiing distance as a constraint according to the designed wing load; s3, determining a second thrust-weight ratio with a single-shot failure climbing gradient as a constraint, a third thrust-weight ratio with a maximum level flight Mach number as a constraint, a fourth thrust-weight ratio with a use lift limit as a constraint and a fifth thrust-weight ratio with an overload as a constraint; and S4, the maximum value in all the thrust-weight ratios is taken as the design thrust-weight ratio of the lower surface blowing flap power high-lift conveyor. According to the method, the design efficiency of the overall parameters of the lower surface blowing type flap power high-lift transport plane is improved.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft design, and in particular relates to a method and device for designing thrust-to-weight ratio of a transport aircraft. Background Art

[0002] The transport wing load is very sensitive to the maximum lift coefficient of the take-off and landing configuration. The maximum lift coefficient increment of the take-off and landing configuration of the blown flap powered lift transport aircraft is closely related to the engine thrust, engine jet efficiency factor, jet deflection angle, and jet wing span-wise influencing factor. The jet efficiency factor and jet deflection angle will also cause a large loss of take-off thrust. The additional trim drag and lift loss caused by engine failure will affect the take-off and climb performance of a single-engine failure. Conventional methods cannot be applied to the design of wing load and thrust-to-weight ratio of blown flap powered lift transport aircraft. Summary of the invention

[0003] In order to solve the above problems, the present application provides a method and device for designing the thrust-to-weight ratio of a transport aircraft, providing technical support for the overall parameter design of a lower surface blown flap powered lift-enhancing transport aircraft.

[0004] In a first aspect, the present application provides a method for designing a thrust-to-weight ratio of a transport aircraft, which is applied to a power-enhanced lift transport aircraft with blown flaps on the lower surface. The method comprises:

[0005] Step S1, calculating the transport aircraft wing loads under different constraint conditions, and taking the minimum value as the design wing load WS of the lower surface blown flap powered lift transport aircraft;

[0006] Step S2: Calculate the first thrust-to-weight ratio TW with the takeoff run distance as a constraint according to the designed wing load WS. 1 ;

[0007] Step S3: Determine the second thrust-to-weight ratio TW with the single engine failure climb gradient as a constraint 2 , the third thrust-to-weight ratio TW with the maximum level flight Mach number as the constraint 3 , the fourth thrust-to-weight ratio TW constrained by the service ceiling 4 and the fifth thrust-to-weight ratio TW with overload as constraint 5 ;

[0008] Step S4, taking the maximum value among the thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

[0009] Preferably, step S1 comprises:

[0010] Step S11: Determine the wing load WS constrained by the cruise design lift coefficient by the following formula: cr :

[0011]

[0012]

[0013]

[0014]

[0015] Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer distance factor, K t is the standby time factor, K z is the comprehensive correction factor, C e is the engine fuel consumption rate.

[0016] Preferably, step S1 comprises:

[0017] Step S12: Determine the wing load WS constrained by the minimum level flight speed of the cruise configuration by the following formula: vm :

[0018]

[0019] Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

[0020] Preferably, step S1 comprises:

[0021] Step S13: Determine the wing load WS constrained by the landing approach speed by the following formula: ap :

[0022]

[0023] C Lmaxland-on =C Lmaxland-off +ΔC Las

[0024] Among them, V ap is the landing approach speed, C Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient, and the jet flap spanwise influence factor φ.

[0025] Preferably, in step S13, the wingspan factor φ affected by the jet is determined by the following formula:

[0026] φ=φ o -φ i ;

[0027] Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

[0028] Preferably, step S1 comprises:

[0029] Step S14: Determine the wing load WS with overload as constraint by the following formula: ny :

[0030]

[0031] Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

[0032] Preferably, in step S2, the first thrust-to-weight ratio TW is determined by the following formula: 1 :

[0033]

[0034] CL maxto-on =CL maxto-off +ΔC Las ;

[0035]

[0036]

[0037] Among them, CL maxto-on is the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80is the percentage of engine thrust loss at a speed of 80 m / s, K h is the altitude correction factor for thrust, ρ 1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

[0038] Preferably, in step S2, the engine jet efficiency factor η is determined by the following formula: j :

[0039]

[0040] Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ 1 =1 and θ d =0, θ jc =0, when λ 1 =1 and θ d ≠0, θ jc =θ d ,on the contrary:

[0041]

[0042]

[0043]

[0044] Among them, θ d is the deflection angle of the engine jet disturbed by the flaps;

[0045] Among them, when F<1.0, When F ≥ 1.0, λ 1 =1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, λ 2 According to the flap downward deflection angle θ relative to the engine axis m Obtained by interpolation calculation.

[0046] Preferably, in step S2, the lower deflection angle θ is determined by the following formula: j :

[0047] θ j =θ jw +θ wi ;

[0048] Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wiThe installation angle of the wing section at the engine installation position.

[0049] Preferably, in step S3, the second thrust-to-weight ratio TW is determined by the following formula: 2 :

[0050]

[0051] Among them, n e is the number of engines, tanθ is the climb gradient, K max The maximum lift-to-drag ratio available for climb with one engine inoperative.

[0052] Preferably, in step S3, the third thrust-to-weight ratio TW is determined by the following formula: 3 :

[0053]

[0054]

[0055] ΔC D =0.00055(M mo -M c )

[0056] K v2 =1-0.32M mo +0.4M mo 2 -0.01M mo 3

[0057] Among them, K v2 is the speed correction factor of thrust, K e K is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number M mo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

[0058] Preferably, in step S3, the fourth thrust-to-weight ratio TW is determined by the following formula: 4 :

[0059]

[0060] K v3 =1-0.32M+0.4M 2 -0.01M 3 ;

[0061] Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max is the maximum lift-to-drag ratio at this speed.

[0062] Preferably, in step S3, the fifth thrust-to-weight ratio TW is determined by the following formula: 5 :

[0063]

[0064] K v4 =1-0.32M 1 +0.4M 1 2 -0.01M 1 3 ;

[0065]

[0066] Among them, q max is the maximum permissible speed, a is the speed of sound, M 1 q for a specified height max The corresponding Mach number, K max is the Mach number M 1 Maximum lift-to-drag ratio under zmax is the maximum normal overload, ρ 2 is the atmospheric density at the current altitude.

[0067] A second aspect of the present application provides a transport aircraft thrust-to-weight ratio design device, which is applied to a lower surface blown flap powered lift transport aircraft, the device comprising:

[0068] The transport aircraft wing load calculation module is used to calculate the transport aircraft wing load under different constraint conditions, and take the minimum value as the design wing load WS of the lower surface blown flap power lift transport aircraft;

[0069] The first thrust-to-weight ratio calculation module is used to calculate the first thrust-to-weight ratio TW based on the design wing load WS with the takeoff run distance as a constraint. 1 ;

[0070] Other thrust-to-weight ratio calculation modules are used to determine the second thrust-to-weight ratio TW with a single engine failure climb gradient as a constraint. 2 , the third thrust-to-weight ratio TW with the maximum level flight Mach number as the constraint 3 , the fourth thrust-to-weight ratio TW constrained by the service ceiling 4 and the fifth thrust-to-weight ratio TW with overload as constraint 5 ;

[0071] The thrust-to-weight ratio selection module is used to take the maximum value among various thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

[0072] Preferably, the transport wing load calculation module includes:

[0073] The first wing load calculation unit is used to determine the wing load WS constrained by the cruise design lift coefficient by the following formula: cr :

[0074]

[0075] q c =0.5ρV c 2 ;

[0076]

[0077]

[0078] Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer distance factor, K t is the standby time factor, K z is the comprehensive correction factor, C e is the engine fuel consumption rate.

[0079] Preferably, the transport wing load calculation module includes:

[0080] The second wing load calculation unit is used to determine the wing load WS constrained by the minimum level flight speed of the cruise configuration through the following formula vm :

[0081]

[0082] Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

[0083] Preferably, the transport wing load calculation module includes:

[0084] The third wing load calculation unit is used to determine the wing load WS constrained by the landing approach speed by the following formula ap :

[0085]

[0086] C Lmaxland-on =C Lmaxland-off +ΔC Las

[0087] Among them, V ap is the landing approach speed, C Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient, and the jet flap spanwise influence factor φ.

[0088] Preferably, in the third wing load calculation unit, the wingspan factor φ affected by the jet is determined by the following formula:

[0089] φ=φ o -φ i ;

[0090] Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

[0091] Preferably, the transport wing load calculation module includes:

[0092] The fourth wing load calculation unit is used to determine the wing load WS with overload as constraint by the following formula ny :

[0093]

[0094] Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

[0095] Preferably, in the first thrust-to-weight ratio calculation module, the first thrust-to-weight ratio TW is determined by the following formula: 1 :

[0096]

[0097] CL maxto-on =CL maxto-off +ΔC Las ;

[0098]

[0099]

[0100] Among them, CL maxto-on is the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80 is the percentage of engine thrust loss at a speed of 80 m / s, K h is the altitude correction factor for thrust, ρ 1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

[0101] Preferably, in the first thrust-to-weight ratio calculation module, the engine jet efficiency factor η is determined by the following formula: j :

[0102]

[0103] Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ 1 =1 and θ d =0, θ jc =0, when λ 1 =1 and θ d ≠0, θ jc =θ d ,on the contrary:

[0104]

[0105]

[0106]

[0107] Among them, θ d is the deflection angle of the engine jet disturbed by the flaps;

[0108] Among them, when F<1.0, When F ≥ 1.0, λ 1=1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, λ 2 According to the flap downward deflection angle θ relative to the engine axis m Obtained by interpolation calculation.

[0109] Preferably, in the first thrust-to-weight ratio calculation module, the lower deflection angle θ is determined by the following formula: j :

[0110] θ j =θ jw +θ wi ;

[0111] Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wi The installation angle of the wing section at the engine installation position.

[0112] Preferably, the other thrust-to-weight ratio calculation module includes a second thrust-to-weight ratio calculation unit, which is used to determine the second thrust-to-weight ratio TW by the following formula: 2 :

[0113]

[0114] Among them, n e is the number of engines, tanθ is the climb gradient, K max The maximum lift-to-drag ratio available for climb with one engine inoperative.

[0115] Preferably, the other thrust-to-weight ratio calculation module includes a third thrust-to-weight ratio calculation unit, which is used to determine the third thrust-to-weight ratio TW by the following formula: 3 :

[0116]

[0117]

[0118] ΔC D =0.00055(M mo -M c )

[0119] K v2 =1-0.32M mo +0.4M mo 2 -0.01M mo 3

[0120] Among them, K v2 is the speed correction factor of thrust, K eK is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number M mo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

[0121] Preferably, the other thrust-to-weight ratio calculation module includes a fourth thrust-to-weight ratio calculation unit, which is used to determine the fourth thrust-to-weight ratio TW by the following formula: 4 :

[0122]

[0123] K v3 =1-0.32M+0.4M 2 -0.01M 3 ;

[0124] Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max is the maximum lift-to-drag ratio at this speed.

[0125] Preferably, the other thrust-to-weight ratio calculation module includes a fifth thrust-to-weight ratio calculation unit, which is used to determine the fifth thrust-to-weight ratio TW by the following formula: 5 :

[0126]

[0127] K v4 =1-0.32M 1 +0.4M 1 2 -0.01M 1 3 ;

[0128]

[0129] Among them, q max is the maximum permissible speed, a is the speed of sound, M 1 q for a specified height max The corresponding Mach number, K max is the Mach number M 1 Maximum lift-to-drag ratio under zmax is the maximum normal overload, ρ 2 is the atmospheric density at the current altitude.

[0130] A third aspect of the present application provides a computer device, comprising a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the method for designing the thrust-to-weight ratio of a transport aircraft as described above.

[0131] A fourth aspect of the present application provides a readable storage medium, wherein the readable storage medium stores a computer program, and when the computer program is executed by a processor, it is used to implement the method for designing the thrust-to-weight ratio of a transport aircraft as described above.

[0132] The present application improves the design efficiency of the overall parameters of a lower surface blown flap powered high-lift transport aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0133] Figure 1 It is a flow chart of a preferred embodiment of the method for designing thrust-to-weight ratio of a transport aircraft of the present application.

[0134] Figure 2 It is a schematic diagram of the structure of a computer device suitable for implementing the embodiments of the present application. DETAILED DESCRIPTION

[0135] In order to make the purpose, technical scheme and advantages of the implementation of this application clearer, the technical scheme in the implementation of this application will be described in more detail in combination with the drawings in the implementation of this application. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and cannot be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in combination with the drawings.

[0136] The first aspect of the present application provides a method for designing the thrust-to-weight ratio of a transport aircraft, which is applied to a power-enhanced lift transport aircraft with blown flaps on the lower surface, such as Figure 1 As shown, the method mainly includes:

[0137] Step S1, calculating the transport aircraft wing loads under different constraint conditions, and taking the minimum value as the design wing load WS of the lower surface blown flap powered lift transport aircraft;

[0138] Step S2: Calculate the first thrust-to-weight ratio TW with the takeoff run distance as a constraint according to the designed wing load WS. 1 ;

[0139] Step S3: Determine the second thrust-to-weight ratio TW with the single engine failure climb gradient as a constraint2 , the third thrust-to-weight ratio TW with the maximum level flight Mach number as the constraint 3 , the fourth thrust-to-weight ratio TW constrained by the service ceiling 4 and the fifth thrust-to-weight ratio TW with overload as constraint 5 ;

[0140] Step S4, taking the maximum value among the thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

[0141] In some optional implementations, step S1 includes:

[0142] Step S11: Determine the wing load WS constrained by the cruise design lift coefficient by the following formula: cr :

[0143]

[0144] q c =0.5ρV c 2 ;

[0145]

[0146]

[0147] Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer range factor, the recommended value is 1.1, K t is the standby time factor, the recommended value is 0.5, K z is the comprehensive correction factor, the recommended value is 1.04, which can be slightly adjusted according to the actual situation. e is the engine fuel consumption rate.

[0148] In some optional implementations, step S1 includes:

[0149] Step S12: Determine the wing load WS constrained by the minimum level flight speed of the cruise configuration by the following formula: vm :

[0150]

[0151] Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

[0152] In some optional implementations, step S1 includes:

[0153] Step S13: Determine the wing load WS constrained by the landing approach speed by the following formula: ap :

[0154]

[0155] C Lmaxland-on =C Lmaxland-off +ΔC Las

[0156] Among them, V ap is the landing approach speed. The maximum landing weight is calculated based on the weight of the aircraft with 70% fuel consumption. Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient Ct, and the jet flap spanwise influence factor φ.

[0157] The following Table 1 shows an interpolation table when φ=0.28.

[0158] Table 1 Data table required for interpolation calculation

[0159]

[0160] In some optional implementations, in step S13, the wingspan factor φ affected by the jet is determined by the following formula:

[0161] φ=φ o -φ i ;

[0162] Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

[0163] In this embodiment, φ o At the spanwise relative position η of the outer boundary line of the outer engine jet at the flap trailing edge o The interpolation calculation is obtained in the first two-dimensional interpolation table composed of ATH. Similarly, φ i At the spanwise relative position η of the inner boundary line of the inner engine jet at the flap trailing edge i It is obtained by interpolation calculation in the second two-dimensional interpolation table composed of ATH. Among them:

[0164]

[0165]

[0166]

[0167] ATH=ARtanΛ 1 / 2 -8λ;

[0168] Among them, X eo , X ei They are the wing span installation positions of the outer and inner engines respectively. jo , D ji are the diameters of the outer and inner engine jets at the trailing edge of the flaps. s is the wing span b w Half of the length, that is, half the span. 1 / 2 is the 1 / 2 chord sweep angle of the wing, and λ is the wing tip-to-root ratio.

[0169] For example, the first two-dimensional interpolation table is shown in Table 2 below.

[0170] Table 2 The first two-dimensional interpolation table

[0171]

[0172] In some optional implementations, step S1 includes:

[0173] Step S14: Determine the wing load WS with overload as constraint by the following formula: ny :

[0174]

[0175] Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

[0176] In some optional implementations, in step S2, the first thrust-to-weight ratio TW is determined by the following formula: 1 :

[0177]

[0178] CL maxto-on =CL maxto-off +ΔC Las ;

[0179]

[0180]

[0181] Among them, CLmaxto-on is the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80 is the percentage of engine thrust loss at a speed of 80 m / s, K h is the altitude correction factor for thrust, ρ 1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

[0182] In some optional embodiments, in step S2, the engine jet efficiency factor η is determined by the following formula: j :

[0183]

[0184] Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ 1 =1 and θ d =0, θ jc =0, when λ 1 =1 and θ d ≠0, θ jc =θ d ,on the contrary:

[0185]

[0186]

[0187]

[0188] Among them, θ d is the deflection angle of the engine jet disturbed by the flaps;

[0189] Among them, when F<1.0, When F ≥ 1.0, λ 1 =1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, λ2 According to the flap downward deflection angle θ relative to the engine axis m Obtained by interpolation calculation.

[0190] In some optional implementations, in step S2, the lower deflection angle θ is determined by the following formula: j :

[0191] θ j =θ jw +θ wi ;

[0192] Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wi The installation angle of the wing section at the engine installation position.

[0193] In this embodiment, θ jw =θ d +θ ew ;

[0194] θ d =asin(sibsinθ m );

[0195] θ m =θ f -θ ew ;

[0196] θ f =δ f +θ fte ;

[0197] θ e =θ ew +θ wi

[0198] Among them, δ f is the flap downward deflection angle, θ fte is the angle between the upper surface of the flap trailing edge and the flap chord, θ ew is the downward deflection angle of the engine axis relative to the wing chord line, θ m is the downward deflection angle of the flap relative to the engine axis, θ d is the deflection angle of the jet disturbed by the flap, and the parameter sib is obtained by interpolation calculation in the given interpolation table.

[0199] In some optional implementations, in step S3, the second thrust-to-weight ratio TW is determined by the following formula: 2 :

[0200]

[0201] Among them, n e is the number of engines, tanθ is the climb gradient, Kmax It is the maximum lift-to-drag ratio available for climb with a single engine failure. It is calculated based on the maximum lift-to-drag ratio for takeoff with all engines, corrected for the trim drag due to single engine failure and deducted from the landing gear resistance.

[0202] In some optional implementations, in step S3, the third thrust-to-weight ratio TW is determined by the following formula: 3 :

[0203]

[0204]

[0205] ΔC D =0.00055(M mo -M c )

[0206] K v2 =1-0.32M mo +0.4M mo 2 -0.01M mo 3

[0207] Among them, K v2 is the speed correction factor of thrust, K e K is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number M mo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

[0208] In some optional implementations, in step S3, the fourth thrust-to-weight ratio TW is determined by the following formula: 4 :

[0209]

[0210] K v3 =1-0.32M+0.4M 2 -0.01M 3 ;

[0211] Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max It is the maximum lift-to-drag ratio at that speed, and the climb weight is calculated by consuming 10% of the mission fuel.

[0212] In some optional implementations, in step S3, the fifth thrust-to-weight ratio TW is determined by the following formula: 5 :

[0213]

[0214] K v4 =1-0.32M 1 +0.4M 1 2 -0.01M 1 3 ;

[0215]

[0216] Among them, q max is the maximum permissible speed, a is the speed of sound, M 1 q for a specified height max The corresponding Mach number, K max is the Mach number M 1 Maximum lift-to-drag ratio under zmax is the maximum normal overload, ρ 2 is the atmospheric density at the current altitude.

[0217] The second aspect of the present application provides a transport aircraft thrust-to-weight ratio design device corresponding to the above method, which is applied to a lower surface blown flap powered lift transport aircraft, and the device comprises:

[0218] The transport aircraft wing load calculation module is used to calculate the transport aircraft wing load under different constraint conditions, and take the minimum value as the design wing load WS of the lower surface blown flap power lift transport aircraft;

[0219] The first thrust-to-weight ratio calculation module is used to calculate the first thrust-to-weight ratio TW based on the design wing load WS with the takeoff run distance as a constraint. 1 ;

[0220] Other thrust-to-weight ratio calculation modules are used to determine the second thrust-to-weight ratio TW with a single engine failure climb gradient as a constraint. 2 , the third thrust-to-weight ratio TW with the maximum level flight Mach number as the constraint 3 , the fourth thrust-to-weight ratio TW constrained by the service ceiling 4 and the fifth thrust-to-weight ratio TW with overload as constraint 5 ;

[0221] The thrust-to-weight ratio selection module is used to take the maximum value among various thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

[0222] In some optional embodiments, the transport wing load calculation module includes:

[0223] The first wing load calculation unit is used to determine the wing load WS constrained by the cruise design lift coefficient by the following formula: cr :

[0224]

[0225] q c =0.5ρV c 2 ;

[0226]

[0227]

[0228] Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer distance factor, K t is the standby time factor, K z is the comprehensive correction factor, C e is the engine fuel consumption rate.

[0229] In some optional embodiments, the transport wing load calculation module includes:

[0230] The second wing load calculation unit is used to determine the wing load WS constrained by the minimum level flight speed of the cruise configuration through the following formula vm :

[0231]

[0232] Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

[0233] In some optional embodiments, the transport wing load calculation module includes:

[0234] The third wing load calculation unit is used to determine the wing load WS constrained by the landing approach speed by the following formula ap :

[0235]

[0236] C Lmaxland-on =C Lmaxland-off +ΔC Las

[0237] Among them, V ap is the landing approach speed, C Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient, and the jet flap spanwise influence factor φ.

[0238] In some optional embodiments, in the third wing load calculation unit, the wingspan factor φ affected by the jet is determined by the following formula:

[0239] φ=φ o -φ i ;

[0240] Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

[0241] In some optional embodiments, the transport wing load calculation module includes:

[0242] The fourth wing load calculation unit is used to determine the wing load WS with overload as constraint by the following formula ny :

[0243]

[0244] Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

[0245] In some optional implementations, in the first thrust-to-weight ratio calculation module, the first thrust-to-weight ratio TW is determined by the following formula: 1 :

[0246]

[0247] CL maxto-on =CL maxto-off +ΔC Las ;

[0248]

[0249]

[0250] Among them, CL maxto-onis the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80 is the percentage of engine thrust loss at a speed of 80 m / s, K h is the altitude correction factor for thrust, ρ 1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

[0251] In some optional embodiments, in the first thrust-to-weight ratio calculation module, the engine jet efficiency factor η is determined by the following formula: j :

[0252]

[0253] Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ 1 =1 and θ d =0, θ jc =0, when λ 1 =1 and θ d ≠0, θ jc =θ d ,on the contrary:

[0254]

[0255]

[0256]

[0257] Among them, θ d is the deflection angle of the engine jet disturbed by the flaps;

[0258] Among them, when F<1.0, When F ≥ 1.0, λ 1 =1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, λ2 According to the flap downward deflection angle θ relative to the engine axis m Obtained by interpolation calculation.

[0259] In some optional implementations, in the first thrust-to-weight ratio calculation module, the downward deflection angle θ is determined by the following formula: j :

[0260] θ j =θ jw +θ wi ;

[0261] Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wi The installation angle of the wing section at the engine installation position.

[0262] In some optional embodiments, the other thrust-to-weight ratio calculation module includes a second thrust-to-weight ratio calculation unit, which is used to determine the second thrust-to-weight ratio TW by the following formula: 2 :

[0263]

[0264] Among them, n e is the number of engines, t an θ is the climbing gradient, K max The maximum lift-to-drag ratio available for climb with one engine inoperative.

[0265] In some optional embodiments, the other thrust-to-weight ratio calculation module includes a third thrust-to-weight ratio calculation unit, which is used to determine the third thrust-to-weight ratio TW by the following formula: 3 :

[0266]

[0267]

[0268] ΔC D =0.00055(M mo -M c )

[0269] K v2 =1-0.32M mo +0.4M mo 2 -0.01M mo 3

[0270] Among them, K v2 is the speed correction factor of thrust, K e K is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number Mmo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

[0271] In some optional embodiments, the other thrust-to-weight ratio calculation module includes a fourth thrust-to-weight ratio calculation unit, which is used to determine the fourth thrust-to-weight ratio TW by the following formula: 4 :

[0272]

[0273] K v3 =1-0.32M+0.4M 2 -0.01M 3 ;

[0274] Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max is the maximum lift-to-drag ratio at this speed.

[0275] In some optional embodiments, the other thrust-to-weight ratio calculation module includes a fifth thrust-to-weight ratio calculation unit, which is used to determine the fifth thrust-to-weight ratio TW by the following formula: 5 :

[0276]

[0277] K v4 =1-0.32M 1 +0.4M 1 2 -0.01M 1 3 ;

[0278]

[0279] Among them, q max is the maximum permissible speed, a is the speed of sound, M 1 q for a specified height max The corresponding Mach number, K max is the Mach number M 1 Maximum lift-to-drag ratio under zmax is the maximum normal overload, ρ 2 is the atmospheric density at the current altitude.

[0280] In a third aspect of the present application, a computer device includes a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement a method for designing a thrust-to-weight ratio of a transport aircraft.

[0281] In a fourth aspect of the present application, a readable storage medium stores a computer program, and when the computer program is executed by a processor, it is used to implement the above-mentioned method for designing the thrust-to-weight ratio of a transport aircraft. The computer-readable storage medium may be included in the device described in the above-mentioned embodiment; or it may exist independently without being assembled into the device. The above-mentioned computer-readable storage medium carries one or more programs, and when the above-mentioned one or more programs are executed by the device, the data is processed according to the above-mentioned method.

[0282] Reference below Figure 2 , which shows a schematic structural diagram of a computer device 400 suitable for implementing the embodiments of the present application. Figure 2 The computer device shown is only an example and should not bring any limitation to the functions and scope of use of the embodiments of the present application.

[0283] like Figure 2 As shown, the computer device 400 includes a central processing unit (CPU) 401, which can perform various appropriate actions and processes according to a program stored in a read-only memory (ROM) 402 or a program loaded from a storage part 408 into a random access memory (RAM) 403. In the RAM 403, various programs and data required for the operation of the device 400 are also stored. The CPU 401, the ROM 402, and the RAM 403 are connected to each other via a bus 404. An input / output (I / O) interface 405 is also connected to the bus 404.

[0284] The following components are connected to the I / O interface 405: an input section 406 including a keyboard, a mouse, etc.; an output section 407 including a cathode ray tube (CRT), a liquid crystal display (LCD), etc., and a speaker, etc.; a storage section 408 including a hard disk, etc.; and a communication section 409 including a network interface card such as a LAN card, a modem, etc. The communication section 409 performs communication processing via a network such as the Internet. A drive 410 is also connected to the I / O interface 405 as needed. A removable medium 411, such as a magnetic disk, an optical disk, a magneto-optical disk, a semiconductor memory, etc., is installed on the drive 410 as needed, so that a computer program read therefrom is installed into the storage section 408 as needed.

[0285] In particular, according to the embodiment of the present application, the process described above with reference to the flowchart can be implemented as a computer software program. For example, the embodiment of the present application includes a computer program product, which includes a computer program carried on a computer-readable medium, and the computer program includes a program code for executing the method shown in the flowchart. In such an embodiment, the computer program can be downloaded and installed from the network through the communication part 409, and / or installed from the removable medium 411. When the computer program is executed by the central processing unit (CPU) 401, the above functions defined in the method of the present application are executed. It should be noted that the computer storage medium of the present application can be a computer-readable signal medium or a computer-readable storage medium or any combination of the above two. The computer-readable storage medium can be, for example, - but not limited to - a system, device or device of electricity, magnetism, light, electromagnetic, infrared, or semiconductor, or any combination of the above. More specific examples of computer-readable storage media may include, but are not limited to, an electrical connection with one or more conductors, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present application, a computer-readable storage medium may be any tangible medium containing or storing a program that can be used by or in combination with an instruction execution system, an apparatus or a device. In the present application, a computer-readable signal medium may include a data signal propagated in a baseband or as part of a carrier wave, which carries a computer-readable program code. Such propagated data signals may take a variety of forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination of the above. A computer-readable signal medium may also be any computer-readable medium other than a computer-readable storage medium, which may send, propagate, or transmit a program for use by or in combination with an instruction execution system, an apparatus or a device. The program code contained on the computer-readable medium may be transmitted using any appropriate medium, including but not limited to: wireless, wire, optical cable, RF, etc., or any suitable combination of the above.

[0286] The flow chart and block diagram in the accompanying drawings illustrate the possible architecture, function and operation of the system, method and computer program product according to various embodiments of the present application. In this regard, each square box in the flow chart or block diagram can represent a module, a program segment or a part of a code, and the module, the program segment or a part of the code contains one or more executable instructions for realizing the specified logical function. It should also be noted that in some implementations as replacements, the functions marked in the square box can also occur in a sequence different from that marked in the accompanying drawings. For example, two square boxes represented in succession can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each square box in the block diagram and / or flow chart, and the combination of the square boxes in the block diagram and / or flow chart can be implemented with a dedicated hardware-based system that performs a specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0287] The modules or units involved in the embodiments described in this application may be implemented by software or hardware. The modules or units described may also be set in a processor, and the names of these modules or units do not constitute limitations on the modules or units themselves in some cases.

[0288] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A method for designing thrust-to-weight ratio of a transport aircraft, applied to a power-lift transport aircraft with blown flaps on the lower surface, characterized in that: The method includes: Step S1, calculating the transport aircraft wing loads under different constraint conditions, and taking the minimum value as the design wing load WS of the lower surface blown flap powered lift transport aircraft; Step S2, calculating the first thrust-to-weight ratio TW1 with the takeoff run distance as a constraint according to the design wing load WS; Step S3, determining a second thrust-to-weight ratio TW2 with a single-engine failure climb gradient as a constraint, a third thrust-to-weight ratio TW3 with a maximum level flight Mach number as a constraint, a fourth thrust-to-weight ratio TW4 with a service ceiling as a constraint, and a fifth thrust-to-weight ratio TW5 with an overload as a constraint; Step S4, taking the maximum value among the thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

2. The method for designing the thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: Step S1 includes: Step S11: Determine the wing load WS constrained by the cruise design lift coefficient by the following formula: cr : Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer distance factor, K t is the standby time factor, K z is the comprehensive correction factor, C e is the engine fuel consumption rate.

3. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: Step S1 includes: Step S12: Determine the wing load WS constrained by the minimum level flight speed of the cruise configuration by the following formula: vm : Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

4. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: Step S1 includes: Step S13: Determine the wing load WS constrained by the landing approach speed by the following formula: ap : Among them, V ap is the landing approach speed, C Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient, and the jet flap spanwise influence factor φ.

5. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 4, characterized in that: In step S13, the wingspan factor φ affected by the jet is determined by the following formula: f=f o -f i ; Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

6. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: Step S1 includes: Step S14: Determine the wing load WS with overload as constraint by the following formula: ny : Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

7. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: In step S2, the first thrust-to-weight ratio TW1 is determined by the following formula: CL maxto-on =CL maxto-off +ΔC Las ; Among them, CL maxto-on is the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80 is the percentage of engine thrust loss at a speed of 80 m / s, K h is the height correction factor for thrust, ρ1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

8. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 7, characterized in that: In step S2, the engine jet efficiency factor η is determined by the following formula: j : Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ1=1 and θ d =0, θ jc =0, when λ1=1 and θ d ≠0, θ jc =θ d ,on the contrary: Among them, θ d is the deflection angle of the engine jet disturbed by the flaps; Among them, when F<1.0, When F ≥ 1.0, λ1 = 1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, and λ2 is determined by the downward deflection angle θ of the flap relative to the engine axis. m Obtained by interpolation calculation.

9. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 7, characterized in that: In step S2, the lower deflection angle θ is determined by the following formula: j : i j =θ jw +θ wi ; Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wi The installation angle of the wing section at the engine installation position.

10. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: In step S3, the second thrust-to-weight ratio TW2 is determined by the following formula: Among them, n e is the number of engines, tanθ is the climb gradient, K max The maximum lift-to-drag ratio available for climb with one engine inoperative.

11. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: In step S3, the third thrust-to-weight ratio TW3 is determined by the following formula: Among them, K v2 is the speed correction factor of thrust, K e K is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number M mo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

12. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: In step S3, the fourth thrust-to-weight ratio TW4 is determined by the following formula: K v3 =1-0.32M+0.4M 2 -0.01M 3 ; Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max is the maximum lift-to-drag ratio at this speed.

13. The method for designing thrust-to-weight ratio of a transport aircraft according to claim 1, characterized in that: In step S3, the fifth thrust-to-weight ratio TW5 is determined by the following formula: K v4 =1-0.32M1+0.4M1 2 -0.01M1 3 ; Among them, q max is the maximum permissible speed, a is the speed of sound, M1 is q at a specified altitude max The corresponding Mach number, K max is the maximum lift-to-drag ratio at Mach number M1, n zmax is the maximum normal overload, and ρ2 is the atmospheric density at the current altitude.

14. A thrust-to-weight ratio design device for a transport aircraft, applied to a power-lift transport aircraft with blown flaps on the lower surface, characterized in that: The device includes: The transport aircraft wing load calculation module is used to calculate the transport aircraft wing load under different constraint conditions, and take the minimum value as the design wing load WS of the lower surface blown flap power lift transport aircraft; A first thrust-to-weight ratio calculation module, used to calculate a first thrust-to-weight ratio TW1 constrained by a takeoff run distance according to a design wing load WS; Other thrust-to-weight ratio calculation modules are used to determine a second thrust-to-weight ratio TW2 constrained by a single-engine failure climb gradient, a third thrust-to-weight ratio TW3 constrained by a maximum level flight Mach number, a fourth thrust-to-weight ratio TW4 constrained by a service ceiling, and a fifth thrust-to-weight ratio TW5 constrained by an overload; The thrust-to-weight ratio selection module is used to take the maximum value among various thrust-to-weight ratios as the design thrust-to-weight ratio of the lower surface blown flap powered lift transport aircraft.

15. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The transport wing load calculation module includes: The first wing load calculation unit is used to determine the wing load WS constrained by the cruise design lift coefficient by the following formula cr : Among them, C Ld Design lift coefficient for cruise, V c is the cruising speed, ρ is the air density at the cruising altitude, L X To design the voyage, W fx is the mission fuel coefficient, K c is the average lift-to-drag ratio for cruising, K R is the transfer distance factor, K t is the standby time factor, K z is the comprehensive correction factor, C e is the engine fuel consumption rate.

16. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The transport wing load calculation module includes: The second wing load calculation unit is used to determine the wing load WS constrained by the minimum level flight speed of the cruise configuration through the following formula vm : Among them, V min is the minimum level flight speed in cruise configuration, C Lmax is the maximum lift coefficient corresponding to the minimum level flight speed in the cruise configuration.

17. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The transport wing load calculation module includes: The third wing load calculation unit is used to determine the wing load WS constrained by the landing approach speed by the following formula ap : Among them, V ap is the landing approach speed, C Lmaxland-on is the maximum lift coefficient of the transport aircraft in the landing configuration in the power-lift state, C Lmaxland-off is the maximum lift coefficient of the unpowered landing configuration, ΔC Las It is the lift coefficient increment caused by the dynamic influence at the stall angle of attack, which is obtained by interpolation calculation in the interpolation table composed of the jet immersion depth factor F of the flap, the thrust coefficient, and the jet flap spanwise influence factor φ.

18. The thrust-to-weight ratio design device for a transport aircraft according to claim 17, characterized in that: In the third wing load calculation unit, the wingspan factor φ affected by the jet is determined by the following formula: f=f o -f i ; Among them, φ o is the spanwise influence factor of the outer engine jet, φ i is the spanwise influence factor of the inner engine jet.

19. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The transport wing load calculation module includes: The fourth wing load calculation unit is used to determine the wing load WS with overload as constraint by the following formula ny : Among them, n zmax is the maximum normal load factor, q max is the maximum allowable velocity pressure, C Lyx is the maximum available lift coefficient.

20. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: In the first thrust-to-weight ratio calculation module, the first thrust-to-weight ratio TW1 is determined by the following formula: CL maxto-on =CL maxto-off +ΔC Las ; Among them, CL maxto-on is the maximum lift coefficient of the takeoff configuration in the power-up state, CL maxto-off is the maximum lift coefficient of the unpowered takeoff configuration, ΔCL as is the stall angle of attack α s The lift coefficient increment caused by dynamic influence, V lof is the take-off speed, L r is the takeoff ground rolling distance, f is the rolling friction coefficient, K ron The lift-to-drag ratio corresponding to the angle of attack α of the takeoff configuration in the power-up state. When the parking angle is zero, α is zero. K v1 is the speed correction factor for takeoff thrust, P x80 is the percentage of engine thrust loss at a speed of 80 m / s, K h is the height correction factor for thrust, ρ1 is the air density at the airport altitude, η j is the engine jet efficiency factor, θ j It is the downward angle of the rear axis of the engine jet deflected by the flap relative to the fuselage axis.

21. The thrust-to-weight ratio design device for a transport aircraft according to claim 20, characterized in that: In the first thrust-to-weight ratio calculation module, the engine jet efficiency factor η is determined by the following formula: j : Among them, θ jc is the theoretical deflection angle of the engine jet caused by flap interference, when λ1=1 and θ d =0, θ jc =0, when λ1=1 and θ d ≠0, θ jc =θ d ,on the contrary: Among them, θ d is the deflection angle of the engine jet disturbed by the flaps; Among them, when F<1.0, When F ≥ 1.0, λ1 = 1; where Y = 1-2F, F is the jet immersion factor of the flap, H f is the jet immersion depth of the flap, and λ2 is determined by the downward deflection angle θ of the flap relative to the engine axis. m Obtained by interpolation calculation.

22. The thrust-to-weight ratio design device for a transport aircraft according to claim 20, characterized in that: In the first thrust-to-weight ratio calculation module, the downward deflection angle θ is determined by the following formula: j : i j =θ jw +θ wi ; Among them, θ jw is the downward deflection angle of the engine jet relative to the wing chord line, θ wi The installation angle of the wing section at the engine installation position.

23. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The other thrust-to-weight ratio calculation module includes a second thrust-to-weight ratio calculation unit, which is used to determine the second thrust-to-weight ratio TW2 by the following formula: Among them, n e is the number of engines, t an θ is the climbing gradient, K max The maximum lift-to-drag ratio available for climb with one engine inoperative.

24. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The other thrust-to-weight ratio calculation module includes a third thrust-to-weight ratio calculation unit, configured to determine a third thrust-to-weight ratio TW3 by the following formula: Among them, K v2 is the speed correction factor of thrust, K e K is the ratio of the maximum rated thrust to the maximum takeoff thrust, mmo is the maximum level flight Mach number M mo The corresponding maximum lift-to-drag ratio, C Lyl is the favorable lift coefficient for cruising, K max is the maximum lift-to-drag ratio at the cruise design point, ΔC D is relative to the cruising speed M c The increase in the drag coefficient.

25. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The other thrust-to-weight ratio calculation module includes a fourth thrust-to-weight ratio calculation unit, configured to determine a fourth thrust-to-weight ratio TW4 by the following formula: K v3 =1-0.32M+0.4M 2 -0.01M 3 ; Among them, V z is the climb rate, V yl For the favorable climbing speed, M is V yl The corresponding Mach number, K max is the maximum lift-to-drag ratio at this speed.

26. The thrust-to-weight ratio design device for a transport aircraft according to claim 14, characterized in that: The other thrust-to-weight ratio calculation module includes a fifth thrust-to-weight ratio calculation unit, configured to determine a fifth thrust-to-weight ratio TW5 by the following formula: K v4 =1-0.32M1+0.4M1 2 -0.01M1 3 ; Among them, q max is the maximum permissible speed, a is the speed of sound, M1 is q at a specified altitude max The corresponding Mach number, K max is the maximum lift-to-drag ratio at Mach number M1, n zmax is the maximum normal overload, and ρ2 is the atmospheric density at the current altitude.

27. A computer device, characterized in that: The method comprises a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the method for designing the thrust-to-weight ratio of a transport aircraft as claimed in any one of claims 1 to 13.

28. A readable storage medium storing a computer program, characterized in that: When the computer program is executed by a processor, it is used to implement the method for designing the thrust-to-weight ratio of a transport aircraft as described in any one of claims 1-13.

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