Comprehensive design method for low-sonic-boom conceptual layout of supersonic civil aircraft

Abstract

The invention provides a supersonic civil aircraft low-sound-boom conceptual layout comprehensive design method, which comprises the following steps of taking low sound boom as a target, and adopting an inverse design method to carry out large-amplitude modification design on aircraft layout to obtain a first aircraft optimization configuration; an afterbody layout analytic design strategy is adopted, optimization design is carried out on the afterbody layout of the first aircraft optimization configuration, the afterbody layout is determined, and therefore a second aircraft optimization configuration is obtained; and adopting a parameterized near-field overpressure distribution method to carry out inverse design on the fuselage of the second aircraft optimization configuration, specifically carrying out shape modification design on the fuselage, so as to obtain a final aircraft low-sonic-boom configuration. According to the method, the rapid design problem of the low-sound-explosion concept layout of the supersonic civil aircraft can be solved, the supersonic civil aircraft concept layout with excellent low-sound-explosion and low-resistance characteristics can be efficiently designed with less computing resource consumption, and the low-sound-explosion and low-resistance performance of a large supersonic civil aircraft is well considered.

Description

Technical Field

[0001] The invention belongs to the technical field of supersonic civil aircraft design, and in particular relates to a comprehensive design method for a low sonic boom conceptual layout of a supersonic civil aircraft. Background Art

[0002] With the rapid economic development, the demand for civil aviation transportation will increase significantly in the future. Speeding up the flight speed and improving the ride comfort are crucial requirements for the future development of civil aircraft. At present, the mainstream civil aircraft is high subsonic civil aircraft. Although the technology is relatively mature, the flight speed is slow. When flying on long-distance routes, especially transoceanic flights, the flight time is too long, and the ride comfort also drops sharply, which makes it difficult to meet the performance requirements of future civil aircraft. Supersonic civil aircraft can greatly improve the related problems of high subsonic civil aircraft and has become one of the main directions for the future development of civil aircraft.

[0003] According to NASA's "N+X" generation supersonic civil aircraft development plan, small supersonic civil aircraft will be developed first in the future, and the ultimate goal is to develop large supersonic civil aircraft. However, the sonic boom intensity is the core issue that restricts the development of future supersonic civil aircraft. Some technical indicators and requirements formulated by the "N+X" generation supersonic civil aircraft development plan are shown in Table 1, which requires that the sonic boom perceived noise level of supersonic civil aircraft in cruise state should not exceed 70PLdB.

[0004] Table 1 “N+X” generation environment and performance indicators

[0005]

[0006] Years of research have shown that reducing the intensity of sonic booms is an extremely challenging problem: every 1 decibel reduction in the ground-perceived noise level of sonic booms means a significant reduction in sound energy. In addition, it is even more difficult to ensure excellent cruise aerodynamic performance while reducing the sonic boom. Since the intensity of the sonic boom is closely related to the weight and size of the aircraft, it is relatively easier to reduce the sonic boom intensity of small aircraft. Therefore, a variety of low-sonic boom and low-drag layouts for small supersonic civil aircraft and supersonic business jets have been proposed in the prior art, such as: US8453961, US6729577, etc. However, for large supersonic civil aircraft, achieving good low-sonic boom and low-drag performance faces more demanding and complex design challenges. Relevant research shows that when the low-sonic boom and low-drag layout of supersonic business jets with excellent performance is applied to large supersonic civil aircraft, it is found that the performance is completely unable to meet the indicators.

[0007] Therefore, with a view to the development of future large supersonic civil aircraft, there is still a major gap in the low sonic boom and low drag layout of large supersonic civil aircraft. Summary of the invention

[0008] In view of the defects of the prior art, the present invention provides a comprehensive design method for the low sonic boom conceptual layout of a supersonic civil aircraft, which can effectively solve the above problems.

[0009] The technical solution adopted by the present invention is as follows:

[0010] The present invention provides a method for comprehensive design of a supersonic civil aircraft low sonic boom concept layout, comprising the following steps:

[0011] Step S1, taking low sonic boom as the goal, adopting an inverse design method to carry out a substantial modification design of the aircraft layout to obtain a first aircraft optimized configuration;

[0012] Step S2, adopting a rear body layout analytical design strategy to optimize the rear body layout of the first aircraft optimized configuration, determine the rear body layout, and thereby obtain a second aircraft optimized configuration;

[0013] Step S3, using a parameterized near-field overpressure distribution method, performs an inverse design on the fuselage of the second aircraft optimized configuration, specifically, performs a modification design on the fuselage, so as to obtain a final aircraft low sonic boom configuration.

[0014] Preferably, step S1 specifically comprises:

[0015] Step S1.1, given aircraft design point parameters, including: aircraft weight W, Mach number Ma, cruising altitude H, aircraft equivalent length L, nose bluntness yf, F function slope k1 and head-to-tail shock wave ratio pf / pr; under the aircraft design point parameters, using the JSGD low sonic boom algorithm, calculate the low sonic boom target F function distribution, and convert the low sonic boom target F function distribution into a target equivalent cross-sectional area distribution;

[0016] Step S1.2, determining the aircraft baseline configuration and design variables;

[0017] Step S1.3, analyzing the equivalent cross-sectional area distribution of the aircraft reference configuration by using a modified linearization method, and comparing it with the target equivalent cross-sectional area distribution in step S1.1, and calculating the square difference difference;

[0018] Step S1.4, determining the optimization objective function and constraint conditions; wherein: the optimization objective function is to minimize the square difference;

[0019] Step S1.5, determining whether the square difference value satisfies the optimization termination condition; if not, executing step S1.6; if satisfied, executing step S1.7;

[0020] Step S1.6, under the constraint conditions, with the minimum square difference as the goal, adjusting the values ​​of the design variables, thereby adjusting the shapes of the fuselage and wings of the aircraft reference configuration, to obtain the adjusted aircraft reference configuration; then returning to step S1.3 for the adjusted aircraft reference configuration, and continuously performing optimization;

[0021] Step S1.7, outputting the aircraft reference configuration adjusted at this time, which is the first aircraft optimized configuration.

[0022] Preferably, the constraint condition is: the upper and lower limits of the fuselage volume required by the aircraft cabin size.

[0023] Preferably, the optimization termination condition is: the square difference obtained in the latest several consecutive iteration cycles no longer decreases and reaches a stable state;

[0024] The adjusting of the shape of the fuselage and wings of the aircraft base configuration specifically includes: adjusting the fuselage shape, wing sweep, wing dihedral angle and wing twist angle of the aircraft base configuration.

[0025] Preferably, step S2 specifically comprises:

[0026] Step S2.1, determining the rear body layout; wherein the rear body layout includes a T-tail, a cross-tail or a V-tail layout;

[0027] Step S2.2, adopting the analytical design strategy of the rear body layout, optimizing the rear body layout of the first aircraft optimized configuration, determining the rear body layout, and thus obtaining the second aircraft optimized configuration.

[0028] Preferably, step S2.2 is specifically as follows: by using the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration, under the constraint of satisfying the following analytical relative position relationship expression, determine the position coordinates of the key points of the tail plane, and then determine the position and plane shape of the tail, and obtain the rear body layout;

[0029]

[0030] in:

[0031] Ma is the Mach number at the design point of the aircraft;

[0032] z 0 is the altitude of the reference position directly below the aircraft;

[0033] x w ,y w and z w are the x, y, and z coordinates of the trailing edge point of the wing tip of the first aircraft in the optimized configuration, respectively;

[0034] xwr ,y wr and z wr are the x, y, and z coordinates of the wing root trailing edge point of the first aircraft optimized configuration, respectively;

[0035] x f ,y f and z f are the x, y, and z coordinates of the fuselage end points of the first aircraft optimized configuration, respectively;

[0036] x hl and z hl are the x and z coordinates of the leading edge point of the tail wing root respectively;

[0037] x hrt and z hrt are the x and z coordinates of the trailing edge point of the tail wing root respectively;

[0038] x htl ,y htl and z htl are the x, y, and z coordinates of the trailing edge of the tail wing tip;

[0039] Thus, the key point position coordinates of the tail plane are determined by the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration; wherein the key points of the wing and fuselage of the first aircraft optimized configuration include the wing tip trailing edge point, the wing root trailing edge point and the fuselage end point; the key points of the tail plane include the tail root leading edge point, the tail root trailing edge point and the tail tip trailing edge point.

[0040] Preferably, in step S3, the second aircraft optimized configuration is used as the initial aircraft shape, and the fuselage is modified and designed by the following method, so as to obtain the final aircraft low sonic boom configuration:

[0041] Step S3.1, at the Mach number Ma of the aircraft design point, the computational fluid dynamics method CFD is used to calculate the constant lift of the current aircraft shape, and the near-field overpressure distribution dp / p at a distance R directly below the aircraft under the design lift coefficient is extracted. ∞ , and the near-field overpressure distribution dp / p is calculated using the following formula ∞ Convert to equivalent cross-sectional area distribution A ec :

[0042]

[0043] Where: γ is the specific heat ratio 1.2; P ∞ is the static pressure of the infinite flow; dp is the difference between the local static pressure and the static pressure of the infinite flow; L is the length of the aircraft along the longitudinal axis x of the fuselage; t is a variable, and the value range of t is 0~L;

[0044] Step S3.2, obtain the near-field overpressure target distribution with the minimum ground perceived noise level PLdB by an optimization method:

[0045] Step S3.2.1: convert the near-field overpressure distribution dp / p obtained in step S3.1 into ∞ Parameterize and select the near-field overpressure distribution dp / p ∞ The maximum and minimum values ​​on the signal waveform are taken as key control points 1 and 2, and they are parameterized into piecewise linear functions; where the coordinates of key control point i are (x i , (dp / p) i ), i=1,2,x i and (dp / p) i Represent the horizontal and vertical coordinates of the key control point i respectively; the horizontal coordinate is the direction along the longitudinal axis of the aircraft, and the vertical coordinate is the vertical direction;

[0046] Step S3.2.2, convert the converted equivalent cross-sectional area distribution A in step S3.1 ec The end value A ec,end As an indicator of the lift of the aircraft, the terminal value A is constrained in the optimization ec,end constant;

[0047] Step S3.2.3, select the starting position x of the aircraft cabin section respectively start , widest position x widest and the end position x end As key station 1, key station 2 and key station 3; at each key station j, j = 1, 2, 3, its conversion equivalent cross-sectional area distribution A ec,j The following two formulas need to be satisfied:

[0048] A ec,upper ≈A ec,j +A eV,upper -A eV,j

[0049] A ec,lower ≈A ec,j +A eV,lower -A eV,j

[0050] Among them: A ec,upper and A ec,lower , are the upper and lower limits of the conversion equivalent cross-sectional area distribution at the key station j; A eV,j A is the volume equivalent cross-sectional area distribution of the current optimized configuration at the key station j; eV,upper and A eV,lower , are the upper and lower limits of the volume equivalent cross-sectional area distribution at the key station j, respectively;

[0051] Step S3.2.4, the horizontal coordinate x of the key control point i in step S3.2.1 i Fixed, in vertical coordinate (dp / p) i As a design variable; near field overpressure distribution dp / p ∞ The generalized Burgers equation is used to propagate to the ground and calculate the ground perceived noise level PLdB. Taking the ground perceived noise level PLdB as the objective function, the ordinate (dp / p) is optimized under the constraints of steps S3.2.2 and S3.2.3. i , so as to minimize the ground perceived noise level PLdB, and the equivalent cross-sectional area distribution corresponding to the obtained optimization solution is the near-field overpressure target distribution;

[0052] Step S3.2.5, convert the near-field overpressure target distribution obtained by optimization in step S3.2.4 into the conversion equivalent cross-sectional area target distribution A ec,T ;

[0053] Step S3.3, adjust the fuselage shape so that the conversion equivalent cross-sectional area distribution of the configuration at the bottom R is consistent with the conversion equivalent cross-sectional area target distribution A ec,T Matching, to obtain the aircraft shape after the inner layer iterative design;

[0054] Step S3.3.1, based on the conversion of the current aircraft shape equivalent cross-sectional area distribution A ec , Volume equivalent cross-sectional area distribution A eV And the conversion equivalent cross-sectional area target distribution A ec,T , according to the following mixed credibility approximation formula, the volume equivalent cross-sectional area target distribution A is obtained: eV,T ;

[0055] A eV,T ≈A eV +A ec,T -A ec

[0056] Step S3.3.2, parameterize the current aircraft shape and analyze the volume equivalent cross-sectional area distribution A of the aircraft shape eV , and calculate the target distribution A of the volume equivalent cross-sectional area eV,T The square difference of the volume equivalent cross-sectional area is obtained;

[0057] Step S3.3.3, taking the minimization of the square difference of the volume equivalent cross-sectional area as the objective function, optimizing the parameters of the aircraft shape until the square difference of the volume equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape;

[0058] Step S3.3.4, perform CFD analysis on the aircraft shape obtained in step S3.3.3 to obtain the near-field overpressure distribution dp / p at position R directly below the aircraft ∞, and converted to equivalent cross-sectional area distribution A ec , and calculate the converted equivalent cross-sectional area distribution A ec And the conversion equivalent cross-sectional area target distribution A ec,T The square difference of the converted equivalent cross-sectional area is obtained;

[0059] Taking the minimization of the square difference of the converted equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the square difference of the converted equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape, which is the aircraft shape after the inner layer iterative design;

[0060] Step S3.4, judging whether the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design, if not, returning to step S3.1, looping through steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design,

[0061] The shape obtained at this time is the final aircraft low sonic boom configuration designed using the comprehensive design method of the low sonic boom concept layout of supersonic civil aircraft.

[0062] The low sonic boom concept layout comprehensive design method for supersonic civil aircraft provided by the present invention has the following advantages:

[0063] The present invention provides a comprehensive design method for a supersonic civil aircraft low sonic boom conceptual layout, in particular, a low sonic boom and low drag aerodynamic layout configuration design method for a large supersonic civil aircraft. The method of the present invention can solve the problem of rapid design of a supersonic civil aircraft low sonic boom conceptual layout, can efficiently design a supersonic civil aircraft conceptual layout with excellent low sonic boom and low drag characteristics with less computing resource consumption, and well takes into account the low sonic boom and low drag performance of a large supersonic civil aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] Figure 1 The present invention provides a flow chart of a comprehensive design method for a supersonic civil aircraft low sonic boom conceptual layout. DETAILED DESCRIPTION

[0065] The following describes in more detail exemplary embodiments of the present disclosure in conjunction with the accompanying drawings. In order to enable a more thorough understanding of the present disclosure and to fully convey the scope of the present disclosure to those skilled in the art, exemplary embodiments of the present disclosure are shown in the accompanying drawings. It should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments described herein. It should be noted that the embodiments of the present invention and the features in the embodiments can be combined with each other without conflict.

[0066] The present invention provides a comprehensive design method for a supersonic civil aircraft low sonic boom conceptual layout, in particular, a low sonic boom and low drag aerodynamic layout configuration design method for a large supersonic civil aircraft. The method of the present invention can solve the problem of rapid design of a supersonic civil aircraft low sonic boom conceptual layout, can efficiently design a supersonic civil aircraft conceptual layout with excellent low sonic boom and low drag characteristics with less computing resource consumption, and well takes into account the low sonic boom and low drag performance of a large supersonic civil aircraft.

[0067] See also Figure 1 The present invention provides a method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft, comprising the following steps:

[0068] Step S1, with the goal of low sonic boom, adopting an inverse design method to carry out a substantial modification design of the aircraft layout to obtain a first aircraft optimized configuration;

[0069] Step S2, adopting a rear body layout analytical design strategy to optimize the rear body layout of the first aircraft optimized configuration, determine the rear body layout, and thus obtain a second aircraft optimized configuration;

[0070] Step S3, using the parameterized near-field overpressure distribution method PNFO, to carry out inverse design on the fuselage of the second aircraft optimized configuration, specifically, to carry out modification design on the fuselage, so as to obtain the final aircraft low sonic boom configuration.

[0071] The following is a detailed description of steps S1 to S3:

[0072] Step S1, with the goal of low sonic boom, adopting an inverse design method to carry out a substantial modification design of the aircraft layout to obtain a first aircraft optimized configuration;

[0073] Step S1 is specifically as follows:

[0074] Step S1.1, given aircraft design point parameters, including: aircraft weight W, Mach number Ma, cruising altitude H, aircraft equivalent length L, nose bluntness yf, F function slope k1 and head-to-tail shock wave ratio pf / pr; under the aircraft design point parameters, using the JSGD low sonic boom algorithm, calculate the low sonic boom target F function distribution, and convert the low sonic boom target F function distribution into a target equivalent cross-sectional area distribution;

[0075] Step S1.2, determining the aircraft baseline configuration and design variables;

[0076] Step S1.3, analyzing the equivalent cross-sectional area distribution of the aircraft reference configuration by using a modified linearization method, and comparing it with the target equivalent cross-sectional area distribution in step S1.1, and calculating the square difference difference;

[0077] Step S1.4, determining the objective function and constraint conditions of optimization; wherein: the objective function of optimization is to minimize the square difference; the constraint conditions are: the upper and lower limits of the fuselage volume required by the aircraft cabin size;

[0078] Step S1.5, determining whether the square difference value satisfies the optimization termination condition; if not, executing step S1.6; if satisfied, executing step S1.7;

[0079] As a specific implementation, the optimization termination condition is: the square difference obtained in the latest several consecutive iteration cycles no longer decreases and reaches a stable state;

[0080] Step S1.6, under the constraint conditions, with the minimum square difference as the goal, adjusting the values ​​of the design variables, thereby adjusting the shapes of the fuselage and wings of the aircraft reference configuration, to obtain the adjusted aircraft reference configuration; then returning to step S1.3 for the adjusted aircraft reference configuration, and continuously performing optimization;

[0081] In this step, the adjustment of the shape of the fuselage and wings of the aircraft reference configuration includes, but is not limited to: adjusting the fuselage shape, wing sweep, wing dihedral angle and wing twist angle of the aircraft reference configuration.

[0082] Step S1.7, outputting the aircraft reference configuration adjusted at this time, which is the first aircraft optimized configuration.

[0083] Step S2, adopting a rear body layout analytical design strategy to optimize the rear body layout of the first aircraft optimized configuration, determine the rear body layout, and thus obtain a second aircraft optimized configuration;

[0084] Step S2 is specifically as follows:

[0085] Step S2.1, determining the rear body layout; wherein the rear body layout includes a T-tail, a cross-tail or a V-tail layout;

[0086] Step S2.2, adopting the analytical design strategy of the rear body layout, optimizing the rear body layout of the first aircraft optimized configuration, determining the rear body layout, and thus obtaining the second aircraft optimized configuration.

[0087] Step S2.2 is specifically as follows:

[0088] By using the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration, under the constraint of satisfying the following analytical relative position relationship expression, the key point position coordinates of the tail plane are determined, and then the position and plane shape of the tail are determined to obtain the rear body layout;

[0089]

[0090] in:

[0091] Ma is the Mach number at the design point of the aircraft;

[0092] z 0 is the altitude of the reference position directly below the aircraft;

[0093] x w ,y w and z w are the x, y, and z coordinates of the trailing edge point of the wing tip of the first aircraft in the optimized configuration, respectively;

[0094] x wr ,y wr and z wr are the x, y, and z coordinates of the wing root trailing edge point of the first aircraft optimized configuration, respectively;

[0095] x f ,y f and z f are the x, y, and z coordinates of the fuselage end points of the first aircraft optimized configuration, respectively;

[0096] x hl and z hl are the x and z coordinates of the leading edge point of the tail wing root respectively;

[0097] x hrt and z hrt are the x and z coordinates of the trailing edge point of the tail wing root respectively;

[0098] x htl ,y htl and z htl are the x, y, and z coordinates of the trailing edge of the tail wing tip;

[0099] Thus, the key point position coordinates of the tail plane are determined by the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration; wherein the key points of the wing and fuselage of the first aircraft optimized configuration include the wing tip trailing edge point, the wing root trailing edge point and the fuselage end point; the key points of the tail plane include the tail root leading edge point, the tail root trailing edge point and the tail tip trailing edge point.

[0100] Step S3, using a parameterized near-field overpressure distribution method, performs an inverse design on the fuselage of the second aircraft optimized configuration, specifically, performs a modification design on the fuselage, so as to obtain a final aircraft low sonic boom configuration.

[0101] In this step, the second aircraft optimized configuration is used as the initial aircraft shape, a parameterized near-field overpressure target distribution that minimizes the ground sonic boom intensity that satisfies the lift constraint and the cabin size constraint is given, and the fuselage shape is inversely designed to obtain the final aircraft low sonic boom configuration. The specific method is:

[0102] Step S3.1, at the Mach number Ma of the aircraft design point, the computational fluid dynamics method CFD is used to calculate the constant lift of the current aircraft shape, and the near-field overpressure distribution dp / p at a distance R directly below the aircraft under the design lift coefficient is extracted. ∞ , and the near-field overpressure distribution dp / p is calculated using the following formula ∞ Convert to equivalent cross-sectional area distribution A ec :

[0103]

[0104] Where: γ is the specific heat ratio 1.2; P ∞ is the static pressure of the infinite flow; dp is the difference between the local static pressure and the static pressure of the infinite flow; L is the length of the aircraft along the longitudinal axis x of the fuselage; t is a variable, and the value range of t is 0~L;

[0105] Step S3.2, obtain the near-field overpressure target distribution with the minimum ground perceived noise level PLdB by an optimization method:

[0106] Step S3.2.1: convert the near-field overpressure distribution dp / p obtained in step S3.1 into ∞ Parameterize and select the near-field overpressure distribution dp / p ∞ The maximum and minimum values ​​on the signal waveform are taken as key control points 1 and 2, and they are parameterized into piecewise linear functions; where the coordinates of key control point i are (x i , (dp / p) i ), i=1,2,x i and (dp / p) i Represent the horizontal and vertical coordinates of the key control point i respectively; the horizontal coordinate is the direction along the longitudinal axis of the aircraft, and the vertical coordinate is the vertical direction;

[0107] Step S3.2.2, convert the converted equivalent cross-sectional area distribution A in step S3.1 ec The end value A ec,end As an indicator of the lift of the aircraft, the terminal value A is constrained in the optimization ec,end constant;

[0108] Step S3.2.3, select the starting position x of the aircraft cabin section respectively start , widest position x widest and the end position x endAs key station 1, key station 2 and key station 3; at each key station j, j = 1, 2, 3, its conversion equivalent cross-sectional area distribution A ec,j The following two formulas need to be satisfied:

[0109] A ec,upper ≈A ec,j +A eV,upper -A eV,j

[0110] A ec,lower ≈A ec,j +A eV,lower -A eV,j

[0111] Among them: A ec,upper and A ec,lower , are the upper and lower limits of the conversion equivalent cross-sectional area distribution at the key station j; A eV,j A is the volume equivalent cross-sectional area distribution of the current optimized configuration at the key station j; eV,upper and A eV,lower , are the upper and lower limits of the volume equivalent cross-sectional area distribution at the key station j, respectively;

[0112] Step S3.2.4, the horizontal coordinate x of the key control point i in step S3.2.1 i Fixed, in vertical coordinate (dp / p) i As a design variable; near field overpressure distribution dp / p ∞ The generalized Burgers equation is used to propagate to the ground and calculate the ground perceived noise level PLdB. Taking the ground perceived noise level PLdB as the objective function, the ordinate (dp / p) is optimized under the constraints of steps S3.2.2 and S3.2.3. i , so as to minimize the ground perceived noise level PLdB, and the equivalent cross-sectional area distribution corresponding to the obtained optimization solution is the near-field overpressure target distribution;

[0113] Step S3.2.5, convert the near-field overpressure target distribution obtained by optimization in step S3.2.4 into the conversion equivalent cross-sectional area target distribution A ec,T ;

[0114] Step S3.3, adjust the fuselage shape so that the conversion equivalent cross-sectional area distribution of the configuration at the bottom R is consistent with the conversion equivalent cross-sectional area target distribution A ec,T Matching, to obtain the aircraft shape after the inner layer iterative design;

[0115] Step S3.3.1, based on the conversion of the current aircraft shape equivalent cross-sectional area distribution A ec , Volume equivalent cross-sectional area distribution A eV And the conversion equivalent cross-sectional area target distribution A ec,T, according to the following mixed credibility approximation formula, the volume equivalent cross-sectional area target distribution A is obtained: eV,T ;

[0116] A eV,T ≈A eV +A ec,T -A ec

[0117] Step S3.3.2, parameterize the current aircraft shape and analyze the volume equivalent cross-sectional area distribution A of the aircraft shape eV , and calculate the target distribution A of the volume equivalent cross-sectional area eV,T The square difference of the volume equivalent cross-sectional area is obtained;

[0118] Step S3.3.3, taking the minimization of the square difference of the volume equivalent cross-sectional area as the objective function, optimizing the parameters of the aircraft shape until the square difference of the volume equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape;

[0119] Step S3.3.4, perform CFD analysis on the aircraft shape obtained in step S3.3.3 to obtain the near-field overpressure distribution dp / p at position R directly below the aircraft ∞ , and converted to equivalent cross-sectional area distribution A ec , and calculate the converted equivalent cross-sectional area distribution A ec And the conversion equivalent cross-sectional area target distribution A ec,T The square difference of the converted equivalent cross-sectional area is obtained;

[0120] Taking the minimization of the square difference of the converted equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the square difference of the converted equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape, which is the aircraft shape after the inner layer iterative design;

[0121] Step S3.4, judging whether the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design, if not, returning to step S3.1, looping through steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design,

[0122] The shape obtained at this time is the final aircraft low sonic boom configuration designed using the comprehensive design method of the low sonic boom concept layout of supersonic civil aircraft.

[0123] It has been verified in practice that the low sonic boom conceptual layout comprehensive design method for supersonic civil aircraft of the present invention is used to optimize the design of the aircraft baseline configuration, and the aircraft low sonic boom configuration is obtained, and its ground sonic boom intensity and drag coefficient are significantly lower than those of the aircraft baseline configuration.

[0124] The invention provides a comprehensive design method for a supersonic civil aircraft low sonic boom concept layout, in particular, a low sonic boom and low drag aerodynamic layout configuration design method for a large supersonic civil aircraft, comprising three steps: using a JSGD inverse design method based on a sonic boom minimization strategy to carry out a substantial modification design of the aircraft layout; using a rear body layout analytical design method to analytically determine the plane shape and position of the rear body layout, such as a T-tail, a cross tail or a V-tail, etc., according to the coordinate geometric relative relationship of key positions of a fuselage and a wing; using a PNFO inverse design method to give a parameterized near-field overpressure target distribution that minimizes the ground sonic boom intensity that meets lift constraints and cabin size constraints, and inversely designing the fuselage shape.

[0125] The method of the present invention can solve the problem of rapid design of the low sonic boom conceptual layout of a supersonic civil aircraft, and can efficiently design a supersonic civil aircraft conceptual layout with excellent low sonic boom and low drag characteristics with less computing resource consumption, thereby giving good consideration to the low sonic boom and low drag performance of large supersonic civil aircraft.

[0126] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principle of the present invention. These improvements and modifications should also be considered as the scope of protection of the present invention.

Claims

1. A comprehensive design method for low sonic boom concept layout of supersonic civil aircraft, characterized in that: The following steps are involved: Step S1, taking low sonic boom as the goal, adopting an inverse design method to carry out a substantial modification design of the aircraft layout to obtain a first aircraft optimized configuration; Step S2, adopting a rear body layout analytical design strategy to optimize the rear body layout of the first aircraft optimized configuration, determine the rear body layout, and thereby obtain a second aircraft optimized configuration; Step S3, using a parameterized near-field overpressure distribution method, performs an inverse design on the fuselage of the second aircraft optimized configuration, specifically, performs a modification design on the fuselage, so as to obtain a final aircraft low sonic boom configuration.

2. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 1, characterized in that: Step S1 is specifically as follows: Step S1.1, given aircraft design point parameters, including: aircraft weight W, Mach number Ma, cruising altitude H, aircraft equivalent length L, nose bluntness yf, F function slope k1 and head-to-tail shock wave ratio pf / pr; under the aircraft design point parameters, using the JSGD low sonic boom algorithm, calculate the low sonic boom target F function distribution, and convert the low sonic boom target F function distribution into a target equivalent cross-sectional area distribution; Step S1.2, determining the aircraft baseline configuration and design variables; Step S1.3, analyzing the equivalent cross-sectional area distribution of the aircraft reference configuration by using a modified linearization method, and comparing it with the target equivalent cross-sectional area distribution in step S1.1, and calculating the square difference difference; Step S1.4, determining the optimization objective function and constraint conditions; wherein: the optimization objective function is to minimize the square difference; Step S1.5, determining whether the square difference value satisfies the optimization termination condition; if not, executing step S1.6; if satisfied, executing step S1.7; Step S1.6, under the constraint conditions, with the minimum square difference as the goal, adjusting the values ​​of the design variables, thereby adjusting the shapes of the fuselage and wings of the aircraft reference configuration, to obtain the adjusted aircraft reference configuration; then returning to step S1.3 for the adjusted aircraft reference configuration, and continuously performing optimization; Step S1.7, outputting the aircraft reference configuration adjusted at this time, which is the first aircraft optimized configuration.

3. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 2 is characterized in that: The constraints are: the upper and lower limits of the fuselage volume required by the aircraft cabin size.

4. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 2 is characterized in that: The optimization termination condition is: the square difference obtained in the last several consecutive iteration cycles no longer decreases and reaches a stable state; The adjusting of the shape of the fuselage and wings of the aircraft base configuration specifically includes: adjusting the fuselage shape, wing sweep, wing dihedral angle and wing twist angle of the aircraft base configuration.

5. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 1, characterized in that: Step S2 is specifically as follows: Step S2.1, determining the rear body layout; wherein the rear body layout includes a T-tail, a cross-tail or a V-tail layout; Step S2.2, adopting the analytical design strategy of the rear body layout, optimizing the rear body layout of the first aircraft optimized configuration, determining the rear body layout, and thus obtaining the second aircraft optimized configuration.

6. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 5 is characterized in that: Step S2.2 is specifically as follows: by using the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration, under the constraint of satisfying the following analytical relative position relationship expression, determine the position coordinates of the key points of the tail plane, and then determine the position and plane shape of the tail, and obtain the rear body layout; in: Ma is the Mach number at the design point of the aircraft; z0 is the altitude of the reference position directly below the aircraft; x w ,y w and z w are the x, y, and z coordinates of the trailing edge point of the wing tip of the first aircraft in the optimized configuration, respectively; x wr ,y wr and z wr are the x, y, and z coordinates of the wing root trailing edge point of the first aircraft optimized configuration, respectively; x f ,y f and z f are the x, y, and z coordinates of the fuselage end points of the first aircraft optimized configuration, respectively; x hl and z hl are the x and z coordinates of the leading edge point of the tail wing root respectively; x hrt and z hrt are the x and z coordinates of the trailing edge point of the tail wing root respectively; x htl ,y htl and z htl are the x, y, and z coordinates of the trailing edge of the tail wing tip; Thus, the key point position coordinates of the tail plane are determined by the geometric coordinates of the key points of the wing and fuselage of the first aircraft optimized configuration; wherein the key points of the wing and fuselage of the first aircraft optimized configuration include the wing tip trailing edge point, the wing root trailing edge point and the fuselage end point; the key points of the tail plane include the tail root leading edge point, the tail root trailing edge point and the tail tip trailing edge point.

7. The method for comprehensive design of low sonic boom concept layout of supersonic civil aircraft according to claim 1, characterized in that: In step S3, the second aircraft optimized configuration is used as the initial aircraft shape, and the fuselage is modified and designed by the following method, so as to obtain the final aircraft low sonic boom configuration: Step S3.1, at the Mach number Ma of the aircraft design point, the computational fluid dynamics method CFD is used to calculate the constant lift of the current aircraft shape, and the near-field overpressure distribution dp / p at a distance R directly below the aircraft under the design lift coefficient is extracted. ∞ , and the near-field overpressure distribution dp / p is calculated using the following formula ∞ Convert to equivalent cross-sectional area distribution A ec : Where: γ is the specific heat ratio 1.2; P ∞ is the static pressure of the infinite flow; dp is the difference between the local static pressure and the static pressure of the infinite flow; L is the length of the aircraft along the longitudinal axis x of the fuselage; t is a variable, and the value range of t is 0~L; Step S3.2, obtain the near-field overpressure target distribution with the minimum ground perceived noise level PLdB by an optimization method: Step S3.2.1: convert the near-field overpressure distribution dp / p obtained in step S3.1 into ∞ Parameterize and select the near-field overpressure distribution dp / p ∞ The maximum and minimum values ​​on the signal waveform are taken as key control points 1 and 2, and they are parameterized into piecewise linear functions; where the coordinates of key control point i are (x i , (dp / p) i ), i=1,2,x i and (dp / p) i Represent the horizontal and vertical coordinates of the key control point i respectively; the horizontal coordinate is the direction along the longitudinal axis of the aircraft, and the vertical coordinate is the vertical direction; Step S3.2.2, convert the converted equivalent cross-sectional area distribution A in step S3.1 ec The end value A ec,end As an indicator of the lift of the aircraft, the terminal value A is constrained in the optimization ec,end constant; Step S3.2.3, select the starting position x of the aircraft cabin section respectively start , widest position x widest and the end position x end As key station 1, key station 2 and key station 3; at each key station j, j = 1, 2, 3, its conversion equivalent cross-sectional area distribution A ec,j The following two formulas need to be satisfied: A ec,upper ≈A ec,j +A eV,upper -A eV,j A ec,lower ≈A ec,j +A eV,lower -A eV,j Among them: A ec,upper and A ec,lower , are the upper and lower limits of the conversion equivalent cross-sectional area distribution at the key station j; A eV,j A is the volume equivalent cross-sectional area distribution of the current optimized configuration at the key station j; eV,upper and A eV,lower , are the upper and lower limits of the volume equivalent cross-sectional area distribution at the key station j, respectively; Step S3.2.4, the horizontal coordinate x of the key control point i in step S3.2.1 i Fixed, in vertical coordinate (dp / p) i As a design variable; near field overpressure distribution dp / p ∞ The generalized Burgers equation is used to propagate to the ground and calculate the ground perceived noise level PLdB. Taking the ground perceived noise level PLdB as the objective function, the ordinate (dp / p) is optimized under the constraints of steps S3.2.2 and S3.2.

3. i , so as to minimize the ground perceived noise level PLdB, and the equivalent cross-sectional area distribution corresponding to the obtained optimization solution is the near-field overpressure target distribution; Step S3.2.5, convert the near-field overpressure target distribution obtained by optimization in step S3.2.4 into the conversion equivalent cross-sectional area target distribution A ec,T ; Step S3.3, adjust the fuselage shape so that the conversion equivalent cross-sectional area distribution of the configuration at the bottom R is consistent with the conversion equivalent cross-sectional area target distribution A ec,T Matching, to obtain the aircraft shape after the inner layer iterative design; Step S3.3.1, based on the conversion of the current aircraft shape equivalent cross-sectional area distribution A ec , Volume equivalent cross-sectional area distribution A eV And the conversion equivalent cross-sectional area target distribution A ec,T , according to the following mixed credibility approximation formula, the volume equivalent cross-sectional area target distribution A is obtained: eV,T ; A eV,T ≈A eV +A ec,T -A ec Step S3.3.2, parameterize the current aircraft shape and analyze the volume equivalent cross-sectional area distribution A of the aircraft shape eV , and calculate the target distribution A of the volume equivalent cross-sectional area eV,T The square difference of the volume equivalent cross-sectional area is obtained; Step S3.3.3, taking the minimization of the square difference of the volume equivalent cross-sectional area as the objective function, optimizing the parameters of the aircraft shape until the square difference of the volume equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape; Step S3.3.4, perform CFD analysis on the aircraft shape obtained in step S3.3.3 to obtain the near-field overpressure distribution dp / p at position R directly below the aircraft ∞ , and converted to equivalent cross-sectional area distribution A ec , and calculate the converted equivalent cross-sectional area distribution A ec And the conversion equivalent cross-sectional area target distribution A ec,T The square difference of the converted equivalent cross-sectional area is obtained; Taking the minimization of the square difference of the converted equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the square difference of the converted equivalent cross-sectional area cannot be further reduced, thereby obtaining an optimized aircraft shape, which is the aircraft shape after the inner layer iterative design; Step S3.4, judging whether the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design, if not, returning to step S3.1, looping through steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner layer iterative design converges to the lift coefficient required by the design, The shape obtained at this time is the final aircraft low sonic boom configuration designed using the comprehensive design method of the low sonic boom concept layout of supersonic civil aircraft.

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