A conceptual layout synthesis design method for low sonic boom of supersonic civil aircraft
By employing reverse design methods and analytical design of the aft body layout, the low-detonation, low-drag layout of a large supersonic civil aircraft was optimized, solving the design challenges of sonic boom intensity and aerodynamic performance, and achieving a highly efficient low-detonation, low-drag flight configuration design.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- NORTHWESTERN POLYTECHNICAL UNIV
- Filing Date
- 2025-01-23
- Publication Date
- 2026-06-19
AI Technical Summary
Existing technologies are insufficient to effectively reduce the sonic boom intensity of large supersonic civil aircraft while maintaining good cruise aerodynamic performance, resulting in unresolved design challenges.
The reverse design method was used to significantly modify the aircraft layout. Combined with the analytical design of the rear body layout and the parameterized near-field overpressure distribution method, the fuselage was optimized to form a low-detonation, low-drag aircraft configuration.
It achieves low sonic boom and low drag performance for large supersonic civil aircraft. With less computational resource consumption, it efficiently designs an aircraft layout with excellent low sonic boom and low drag characteristics, taking into account both sonic boom and drag performance.
Smart Images
Figure CN120024504B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of supersonic civil aircraft design technology, specifically relating to a comprehensive design method for the low-detonation conceptual layout of a supersonic civil aircraft. Background Technology
[0002] With rapid economic development, the demand for civil aviation transportation will increase significantly in the future. Accelerating flight speed and improving passenger comfort are crucial requirements for the future development of civil aircraft. Currently, the mainstream civil aircraft are high-subsonic aircraft. Although the technology is relatively mature, their flight speed is relatively slow. On long-haul routes, especially transoceanic flights, the flight time is too long, and passenger comfort drops sharply, making it difficult to meet the performance requirements of future civil aircraft. Supersonic aircraft can greatly improve the related problems of high-subsonic aircraft and have become one of the main directions for the future development of civil aircraft.
[0003] According to NASA's "N+X" generation supersonic civilian aircraft development plan, the initial focus will be on developing small supersonic civilian aircraft, with the ultimate goal of developing large supersonic civilian aircraft. However, sonic boom intensity is a core issue restricting the future development of supersonic civilian aircraft. Table 1 shows some of the technical requirements set forth in the "N+X" generation supersonic civilian aircraft development plan, which requires that the perceived sonic boom noise level of a supersonic civilian aircraft during cruise should not exceed 70 PLdB.
[0004] Table 1. Environmental and Performance Indicators for “N+X”
[0005]
[0006] Years of research have shown that reducing sonic boom intensity is an extremely challenging problem: every 1 decibel reduction in the perceived ground noise level of a sonic boom signifies a significant reduction in sound energy. Furthermore, maintaining excellent cruise aerodynamic performance while reducing sonic boom is even more difficult. Since sonic boom intensity is closely related to the weight and size of the aircraft, reducing sonic boom intensity is relatively easier for smaller aircraft. Therefore, various low-sonic-boom, low-drag configurations for small supersonic civil aircraft and supersonic business jets have been proposed in existing technologies, such as US8453961 and US6729577. However, for large supersonic civil aircraft, achieving good low-sonic-boom, low-drag performance presents more stringent and complex design challenges. Related research indicates that applying the high-performance low-sonic-boom, low-drag configurations of supersonic business jets to large supersonic civil aircraft reveals that the performance completely fails to meet the specifications.
[0007] Therefore, there is still a significant gap in the development of future large supersonic civil aircraft, particularly in the design of low-detonation, low-drag configurations for such aircraft. Summary of the Invention
[0008] To address the shortcomings of existing technologies, this invention provides a comprehensive design method for the low-detonation conceptual layout of supersonic civil aircraft, which can effectively solve the above-mentioned problems.
[0009] The technical solution adopted in this invention is as follows:
[0010] This invention provides a comprehensive design method for the low-sonic-boom conceptual layout of a supersonic civil aircraft, comprising the following steps:
[0011] Step S1: With low sonic boom as the target, the reverse design method is used to carry out a major modification design of the aircraft layout to obtain the first optimized aircraft configuration.
[0012] Step S2: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration.
[0013] Step S3: Using the parametric near-field overpressure distribution method, reverse design is carried out on the fuselage of the second optimized aircraft configuration. Specifically, the fuselage is modified to obtain the final low-sonic-bang configuration of the aircraft.
[0014] Preferably, step S1 specifically includes:
[0015] Step S1.1: Given the 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 nose-to-tail shock ratio pf / pr; under the aircraft design point parameters, the JSGD low-sonic blast algorithm is used to calculate the low-sonic blast target F-function distribution, and the low-sonic blast target F-function distribution is converted into the target equivalent cross-sectional area distribution;
[0016] Step S1.2: Determine the aircraft baseline configuration and design variables;
[0017] Step S1.3: The equivalent cross-sectional area distribution of the aircraft reference configuration is analyzed using the modified linearization method, and compared with the target equivalent cross-sectional area distribution in step S1.1 to calculate the difference in squared area.
[0018] Step S1.4: Determine the objective function and constraints for optimization; wherein: the objective function for optimization is to minimize the difference in squared differences;
[0019] Step S1.5: Determine whether the squared difference value meets the optimization termination condition; if not, proceed to step S1.6; if it does, proceed to step S1.7.
[0020] Step S1.6: Under the constraints, with the goal of minimizing the difference in squares, adjust the values of the design variables to adjust the shape of the fuselage and wings of the aircraft reference configuration, thereby obtaining the adjusted aircraft reference configuration; then, for the adjusted aircraft reference configuration, return to step S1.3, and continue to optimize in this way;
[0021] Step S1.7: Output the adjusted aircraft baseline configuration, which is the first optimized aircraft configuration.
[0022] Preferably, the constraint is the upper and lower limits of the fuselage volume required by the aircraft cabin size.
[0023] Preferably, the optimization termination condition is: the difference in squared differences obtained in the most recent consecutive iterations no longer decreases, reaching a stable state;
[0024] The adjustment of the fuselage and wing shape of the aircraft's baseline configuration specifically involves adjusting the fuselage shape, wing sweep, wing dihedral angle, and wing twist angle of the aircraft's baseline configuration.
[0025] Preferably, step S2 specifically includes:
[0026] Step S2.1, determine 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: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration.
[0028] Preferably, step S2.2 specifically involves: using the geometric coordinates of the key points of the wings and fuselage of the first aircraft optimized configuration, under the constraint of satisfying the following analytical relative position relationship expression, determining the position coordinates of the key points of the tail plane, and then determining the position and planar shape of the tail to obtain the rear body layout;
[0029]
[0030] in:
[0031] Ma is the Mach number at the aircraft design point;
[0032] z0 is the altitude of the reference position directly below the aircraft;
[0033] x w y w and z w These are the x, y, and z coordinates of the trailing edge point of the wingtip of the first aircraft's optimized configuration;
[0034] x wry wr and z wr These are the x, y, and z coordinates of the trailing edge point at the wing root of the first aircraft's optimized configuration;
[0035] x f y f and z f These are the x, y, and z coordinates of the fuselage end point of the first aircraft's optimized configuration;
[0036] x hl and z hl These are the x and z coordinates of the leading edge point at the tail fin root;
[0037] x hrt and z hrt These are the x and z coordinates of the trailing edge point at the tail fin root, respectively.
[0038] x htl y htl and z htl These are the x, y, and z coordinates of the trailing edge point of the tail fin tip;
[0039] Therefore, the key point coordinates of the tail plane are determined by using 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 trailing edge point of the wingtip, the trailing edge point of the wing root, and the fuselage tip point; the key points of the tail plane include the leading edge point of the tail root, the trailing edge point of the tail root, and the trailing edge point of the tailtip.
[0040] Preferably, in step S3, using the second optimized aircraft configuration as the initial aircraft shape, the fuselage is modified using the following method to obtain the final low-detonation aircraft configuration:
[0041] Step S3.1: At the Mach number Ma at the aircraft design point, use computational fluid dynamics (CFD) to calculate the constant lift of the current aircraft shape, and extract the near-field overpressure distribution dp / p at a distance R directly below the aircraft under the design lift coefficient. ∞ The following formula is used to determine the near-field overpressure distribution dp / p. ∞ Converted to equivalent cross-sectional area distribution A ec :
[0042]
[0043] Where: γ is the specific heat ratio of 1.2; P ∞ dp is the static pressure of the infinite incoming flow; L is the difference between the local static pressure and the static pressure of the infinite incoming flow; t is the length of the aircraft along the longitudinal axis x; t is a variable, and the value of t ranges from 0 to L.
[0044] Step S3.2: Obtain the near-field overpressure target distribution with the minimum ground perceived noise level (PLdB) using an optimization method.
[0045] Step S3.2.1, the near-field overpressure distribution dp / p obtained in step S3.1 is... ∞ Parameterization was performed, and the near-field overpressure distribution dp / p was selected. ∞ The maximum and minimum values on the signal waveform are used as key control points 1 and 2, respectively, and parameterized as 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 These represent the x-coordinate and y-coordinate of the critical control point i, respectively; the x-coordinate is along the longitudinal axis of the aircraft, and the y-coordinate is in the vertical direction.
[0046] Step S3.2.2, transform the equivalent cross-sectional area distribution A from step S3.1. ec The terminal value A ec,end As an indicator of aircraft lift, 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 stations 1, 2, and 3; at each key station j, j = 1, 2, 3, its transformed equivalent cross-sectional area distribution A ec,j The following two equations must 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 , respectively, are the upper and lower limits of the equivalent cross-sectional area distribution at the key station j; A eV,j The current optimized configuration has an equivalent cross-sectional area distribution at the key station j; A eV,upper and A eV,lower , which are the upper and lower limits of the volume equivalent cross-sectional area distribution at the key station j, respectively;
[0051] The x-coordinate of the key control point i in step S3.2.4 and step S3.2.1 is... i Fixed, with the ordinate (dp / p) i As a design variable; near-field overpressure distribution dp / p ∞ The generalized Burgers equation is propagated to the ground, and the perceived ground noise level PLdB is calculated. Using the perceived ground 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 To minimize the ground-perceived noise level PLdB, the equivalent cross-sectional area distribution corresponding to the optimized solution is the near-field overpressure target distribution.
[0052] Step S3.2.5: Convert the near-field overpressure target distribution obtained from the optimization in step S3.2.4 into a converted equivalent cross-sectional area target distribution A. ec,T ;
[0053] Step S3.3: Adjust the fuselage shape so that the converted equivalent cross-sectional area distribution at point R directly below the fuselage matches the target converted equivalent cross-sectional area distribution A. ec,T Matching is performed to obtain the aircraft shape after the inner iterative design;
[0054] Step S3.3.1, based on the current aircraft shape transformation equivalent cross-sectional area distribution A ec Volume equivalent cross-sectional area distribution A eV and the conversion of equivalent cross-sectional area target distribution A ec,T The volume equivalent cross-sectional area target distribution A is obtained according to the following mixed confidence approximation formula. 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 with volume equivalent cross-sectional area. eV,T The difference of squares is used to obtain the difference of squares of the equivalent cross-sectional area of the volume;
[0057] Step S3.3.3: Using the minimization of the difference in the square of the equivalent cross-sectional area of the volume as the objective function, optimize the parameters of the aircraft shape until the difference in the square of the equivalent cross-sectional area of the volume cannot be further reduced, and obtain the 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 convert it into the equivalent cross-sectional area distribution A. ec And the equivalent cross-sectional area distribution A of the transformation is calculated. ec and the conversion of equivalent cross-sectional area target distribution A ec,T The squared difference is used to obtain the squared difference of the equivalent cross-sectional area.
[0059] Using the minimization of the squared difference of the equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the squared difference of the equivalent cross-sectional area cannot be further reduced, thus obtaining the optimized aircraft shape, which is the aircraft shape after the inner iterative design.
[0060] Step S3.4: Determine whether the lift coefficient of the aircraft shape after the inner iterative design has converged to the required lift coefficient. If not, return to step S3.1 and repeat steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner iterative design converges to the required lift coefficient.
[0061] The shape obtained at this point is the final low-sonic-boom configuration of the aircraft designed using the integrated design method of the supersonic civil aircraft low-sonic-boom concept layout.
[0062] The integrated design method for low-sonic-boom conceptual layout of a supersonic civil aircraft provided by this invention has the following advantages:
[0063] This invention provides a comprehensive design method for the low-detonation conceptual layout of supersonic civil aircraft, particularly a design method for the low-detonation, low-drag aerodynamic layout configuration of large supersonic civil aircraft. Using this method, the rapid design challenge of the low-detonation conceptual layout of supersonic civil aircraft can be solved. It can efficiently design a supersonic civil aircraft conceptual layout with excellent low-detonation and low-drag characteristics with less computational resource consumption, and well balance the low-detonation and low-drag performance of large supersonic civil aircraft. Attached Figure Description
[0064] Figure 1 The flowchart illustrates a comprehensive design method for the low-sonic-bang conceptual layout of a supersonic civil aircraft, as provided by this invention. Detailed Implementation
[0065] Exemplary embodiments of the present disclosure are described in more detail below with reference to the accompanying drawings. Exemplary embodiments of the present disclosure are shown in the accompanying drawings to provide a more thorough understanding of the present disclosure and to fully convey its scope to those skilled in the art. It should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments described herein. It should be noted that, without conflict, the embodiments and features described herein can be combined with each other.
[0066] This invention provides a comprehensive design method for the low-detonation conceptual layout of supersonic civil aircraft, particularly a design method for the low-detonation, low-drag aerodynamic layout configuration of large supersonic civil aircraft. Using this method, the rapid design challenge of the low-detonation conceptual layout of supersonic civil aircraft can be solved. It can efficiently design a supersonic civil aircraft conceptual layout with excellent low-detonation and low-drag characteristics with less computational resource consumption, and well balance the low-detonation and low-drag performance of large supersonic civil aircraft.
[0067] See Figure 1 This invention provides a comprehensive design method for the low-sonic-boom concept layout of a supersonic civil aircraft, comprising the following steps:
[0068] Step S1: With low sonic boom as the target, the reverse design method is used to carry out a major modification design of the aircraft layout to obtain the first optimized aircraft configuration.
[0069] Step S2: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration.
[0070] Step S3: Using the parametric near-field overpressure distribution method PNFO, reverse design is carried out on the fuselage of the second optimized aircraft configuration. Specifically, the fuselage is modified to obtain the final low-sonic-bang configuration of the aircraft.
[0071] The following is a detailed description of steps S1 to S3:
[0072] Step S1: With low sonic boom as the target, the reverse design method is used to carry out a major modification design of the aircraft layout to obtain the first optimized aircraft configuration.
[0073] Step S1 is as follows:
[0074] Step S1.1: Given the 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 nose-to-tail shock ratio pf / pr; under the aircraft design point parameters, the JSGD low-sonic blast algorithm is used to calculate the low-sonic blast target F-function distribution, and the low-sonic blast target F-function distribution is converted into the target equivalent cross-sectional area distribution;
[0075] Step S1.2: Determine the aircraft baseline configuration and design variables;
[0076] Step S1.3: The equivalent cross-sectional area distribution of the aircraft reference configuration is analyzed using the modified linearization method, and compared with the target equivalent cross-sectional area distribution in step S1.1 to calculate the difference in squared area.
[0077] Step S1.4: Determine the objective function and constraints for optimization; wherein: the objective function for optimization is to minimize the difference in squared differences; the constraints are: the upper and lower limits of the fuselage volume required by the aircraft cabin size;
[0078] Step S1.5: Determine whether the squared difference value meets the optimization termination condition; if not, proceed to step S1.6; if it does, proceed to step S1.7.
[0079] As one specific implementation, the optimization termination condition is: the difference in squared differences obtained in the most recent consecutive iterations no longer decreases, reaching a stable state;
[0080] Step S1.6: Under the constraints, with the goal of minimizing the difference in squares, adjust the values of the design variables to adjust the shape of the fuselage and wings of the aircraft reference configuration, thereby obtaining the adjusted aircraft reference configuration; then, for the adjusted aircraft reference configuration, return to step S1.3, and continue to optimize in this way;
[0081] In this step, adjusting the shape of the fuselage and wings of the aircraft's baseline configuration includes, but is not limited to: adjusting the fuselage shape, wing sweep, wing dihedral angle, and wing twist angle of the aircraft's baseline configuration.
[0082] Step S1.7: Output the adjusted aircraft baseline configuration, which is the first optimized aircraft configuration.
[0083] Step S2: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration.
[0084] Step S2 is as follows:
[0085] Step S2.1, determine 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: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration.
[0087] Step S2.2 specifically includes:
[0088] By using the geometric coordinates of the key points of the wings and fuselage of the first aircraft optimized configuration, and under the constraint of the following analytical relative position relationship expression, the position coordinates of the key points of the tail plane are determined, and then the position and planar shape of the tail are determined, thus obtaining the rear body layout.
[0089]
[0090] in:
[0091] Ma is the Mach number at the aircraft design point;
[0092] z0 is the altitude of the reference position directly below the aircraft;
[0093] x w y w and z w These are the x, y, and z coordinates of the trailing edge point of the wingtip of the first aircraft's optimized configuration;
[0094] x wr y wr and z wr These are the x, y, and z coordinates of the trailing edge point at the wing root of the first aircraft's optimized configuration;
[0095] x f y f and z f These are the x, y, and z coordinates of the fuselage end point of the first aircraft's optimized configuration;
[0096] x hl and z hl These are the x and z coordinates of the leading edge point at the tail fin root;
[0097] x hrt and z hrt These are the x and z coordinates of the trailing edge point at the tail fin root, respectively.
[0098] x htl y htl and z htl These are the x, y, and z coordinates of the trailing edge point of the tail fin tip;
[0099] Therefore, the key point coordinates of the tail plane are determined by using 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 trailing edge point of the wingtip, the trailing edge point of the wing root, and the fuselage tip point; the key points of the tail plane include the leading edge point of the tail root, the trailing edge point of the tail root, and the trailing edge point of the tailtip.
[0100] Step S3: Using the parametric near-field overpressure distribution method, reverse design is carried out on the fuselage of the second optimized aircraft configuration. Specifically, the fuselage is modified to obtain the final low-sonic-bang configuration of the aircraft.
[0101] In this step, the second optimized aircraft configuration is used as the initial aircraft shape. A parameterized near-field overpressure target distribution that minimizes ground sonic boom intensity while satisfying lift constraints and cabin size constraints is given. The fuselage shape is then reverse-designed to obtain the final low-sonic-boom aircraft configuration. The specific method is as follows:
[0102] Step S3.1: At the Mach number Ma at the aircraft design point, use computational fluid dynamics (CFD) to calculate the constant lift of the current aircraft shape, and extract the near-field overpressure distribution dp / p at a distance R directly below the aircraft under the design lift coefficient. ∞ The following formula is used to determine the near-field overpressure distribution dp / p. ∞ Converted to equivalent cross-sectional area distribution A ec :
[0103]
[0104] Where: γ is the specific heat ratio of 1.2; P ∞ dp is the static pressure of the infinite incoming flow; L is the difference between the local static pressure and the static pressure of the infinite incoming flow; t is the length of the aircraft along the longitudinal axis x; t is a variable, and the value of t ranges from 0 to L.
[0105] Step S3.2: Obtain the near-field overpressure target distribution with the minimum ground perceived noise level (PLdB) using an optimization method.
[0106] Step S3.2.1, the near-field overpressure distribution dp / p obtained in step S3.1 is... ∞ Parameterization was performed, and the near-field overpressure distribution dp / p was selected. ∞ The maximum and minimum values on the signal waveform are used as key control points 1 and 2, respectively, and parameterized as 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 These represent the x-coordinate and y-coordinate of the critical control point i, respectively; the x-coordinate is along the longitudinal axis of the aircraft, and the y-coordinate is in the vertical direction.
[0107] Step S3.2.2, transform the equivalent cross-sectional area distribution A from step S3.1. ec The terminal value A ec,end As an indicator of aircraft lift, 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 end As key stations 1, 2, and 3; at each key station j, j = 1, 2, 3, its transformed equivalent cross-sectional area distribution A ec,j The following two equations must be satisfied:
[0109] Aec,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 , respectively, are the upper and lower limits of the equivalent cross-sectional area distribution at the key station j; A eV,j The current optimized configuration has an equivalent cross-sectional area distribution at the key station j; A eV,upper and A eV,lower , which are the upper and lower limits of the volume equivalent cross-sectional area distribution at the key station j, respectively;
[0112] The x-coordinate of the key control point i in step S3.2.4 and step S3.2.1 is... i Fixed, with the ordinate (dp / p) i As a design variable; near-field overpressure distribution dp / p ∞ The generalized Burgers equation is propagated to the ground, and the perceived ground noise level PLdB is calculated. Using the perceived ground 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 To minimize the ground-perceived noise level PLdB, the equivalent cross-sectional area distribution corresponding to the optimized solution is the near-field overpressure target distribution.
[0113] Step S3.2.5: Convert the near-field overpressure target distribution obtained from the optimization in step S3.2.4 into a converted equivalent cross-sectional area target distribution A. ec,T ;
[0114] Step S3.3: Adjust the fuselage shape so that the converted equivalent cross-sectional area distribution at point R directly below the fuselage matches the target converted equivalent cross-sectional area distribution A. ec,T Matching is performed to obtain the aircraft shape after the inner iterative design;
[0115] Step S3.3.1, based on the current aircraft shape transformation equivalent cross-sectional area distribution A ec Volume equivalent cross-sectional area distribution A eV and the conversion of equivalent cross-sectional area target distribution A ec,T The volume equivalent cross-sectional area target distribution A is obtained according to the following mixed confidence approximation formula. eV,T ;
[0116] A eV,T ≈A eV +Aec,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 with volume equivalent cross-sectional area. eV,T The difference of squares is used to obtain the difference of squares of the equivalent cross-sectional area of the volume;
[0118] Step S3.3.3: Using the minimization of the difference in the square of the equivalent cross-sectional area of the volume as the objective function, optimize the parameters of the aircraft shape until the difference in the square of the equivalent cross-sectional area of the volume cannot be further reduced, and obtain the 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 convert it into the equivalent cross-sectional area distribution A. ec And the equivalent cross-sectional area distribution A of the transformation is calculated. ec and the conversion of equivalent cross-sectional area target distribution A ec,T The squared difference is used to obtain the squared difference of the equivalent cross-sectional area.
[0120] Using the minimization of the squared difference of the equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the squared difference of the equivalent cross-sectional area cannot be further reduced, thus obtaining the optimized aircraft shape, which is the aircraft shape after the inner iterative design.
[0121] Step S3.4: Determine whether the lift coefficient of the aircraft shape after the inner iterative design has converged to the required lift coefficient. If not, return to step S3.1 and repeat steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner iterative design converges to the required lift coefficient.
[0122] The shape obtained at this point is the final low-sonic-boom configuration of the aircraft designed using the integrated design method of the supersonic civil aircraft low-sonic-boom concept layout.
[0123] Practical verification has shown that by adopting the integrated design method of the supersonic civil aircraft low sonic boom concept layout of the present invention, the basic configuration of the aircraft is optimized and a low sonic boom configuration is obtained, whose ground sonic boom intensity and drag coefficient are significantly lower than those of the basic configuration of the aircraft.
[0124] This invention provides a comprehensive design method for the low-detonation conceptual layout of supersonic civil aircraft, particularly a design method for the low-detonation, low-drag aerodynamic layout configuration of large supersonic civil aircraft. The method comprises three steps: First, a significant modification design of the aircraft layout is performed using the JSGD inverse design method based on a sonic boom minimization strategy. Second, an analytical design method for the aft body layout is used to analytically determine the planar shape and position of the aft body layout, such as a T-tail, cruciform tail, or V-tail, based on the relative coordinate geometry of key positions on the fuselage and wings. Third, the PNFO inverse design method is used to provide a parameterized near-field overpressure target distribution that minimizes ground sonic boom intensity while satisfying lift and cabin size constraints, and to perform inverse design on the fuselage shape.
[0125] The method of this invention can solve the problem of rapid design of the low-detonation concept layout of supersonic civil aircraft. It can efficiently design a supersonic civil aircraft concept layout with excellent low-detonation and low-drag characteristics with less computational resource consumption, and well balance the low-detonation and low-drag performance of large supersonic civil aircraft.
[0126] The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.
Claims
1. A comprehensive design method for the low-sonic-boom conceptual layout of a supersonic civil aircraft, characterized in that, Includes the following steps: Step S1: With low sonic boom as the target, the reverse design method is used to carry out a major modification design of the aircraft layout to obtain the first optimized aircraft configuration. Step S2: Using the aft body layout analytical design strategy, the aft body layout of the first aircraft optimized configuration is optimized to determine the aft body layout, thereby obtaining the second aircraft optimized configuration. Step S3: Using the parametric near-field overpressure distribution method, reverse design is carried out on the fuselage of the second optimized aircraft configuration. Specifically, the fuselage is modified to obtain the final low-sonic-bang configuration of the aircraft. Step S1 is as follows: Step S1.1: Given the 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 nose-to-tail shock ratio pf / pr; under the aircraft design point parameters, the JSGD low-sonic blast algorithm is used to calculate the low-sonic blast target F-function distribution, and the low-sonic blast target F-function distribution is converted into the target equivalent cross-sectional area distribution; Step S1.2: Determine the aircraft baseline configuration and design variables; Step S1.3: The equivalent cross-sectional area distribution of the aircraft reference configuration is analyzed using the modified linearization method, and compared with the target equivalent cross-sectional area distribution in step S1.1 to calculate the difference in squared area. Step S1.4: Determine the objective function and constraints for optimization; wherein: the objective function for optimization is to minimize the difference in squared differences; the constraints are: the upper and lower limits of the fuselage volume required by the aircraft cabin dimensions; Step S1.5: Determine whether the squared difference value meets the optimization termination condition; if not, proceed to step S1.6; if it does, proceed to step S1.7; the optimization termination condition is: the squared difference value obtained in the most recent consecutive iterations no longer decreases, reaching a stable state; Step S1.6: Under the constraints, with the goal of minimizing the difference in squares, adjust the values of the design variables to adjust the shape of the fuselage and wings of the aircraft reference configuration, thereby obtaining the adjusted aircraft reference configuration; then, for the adjusted aircraft reference configuration, return to step S1.3, and continue to optimize in this way; the adjustment of the shape of the fuselage and wings of the aircraft reference configuration specifically involves: adjusting the fuselage shape, wing sweep, wing dihedral angle, and wing twist angle of the aircraft reference configuration; Step S1.7: Output the adjusted aircraft baseline configuration, which is the first optimized aircraft configuration. Step S2 is as follows: Step S2.1, determine the rear body layout; wherein, the rear body layout includes a T-tail, a cross-tail, or a V-tail layout; Step S2.2: Using the aft body layout analytical design strategy, optimize the aft body layout of the first aircraft optimized configuration, determine the aft body layout, and thus obtain the second aircraft optimized configuration. Step S2.2 specifically involves: using the geometric coordinates of the key points of the wings and fuselage of the first aircraft optimized configuration, and under the constraint of satisfying the following analytical relative position relationship expression, determining the position coordinates of the key points of the tail plane, thereby determining the position and planar shape of the tail, and obtaining the rear body layout; ； in: Ma is the Mach number at the aircraft design point; z0 is the altitude of the reference position directly below the aircraft; x w , y w , and z w are the x, y, z coordinates, respectively, of the wing tip trailing edge point of the first aircraft optimized configuration; x wr , y wr , and z wr are the x, y, z coordinates, respectively, of the wing root trailing edge point of the first aircraft optimized configuration; x f , y f , and z f are the x, y, z coordinates of the end of the fuselage of the first aircraft optimized configuration, respectively; x hl and z hl are the x, z coordinates of the tail wing root leading edge point, respectively; x hrt and z hrt These are the x and z coordinates of the trailing edge point at the tail fin root, respectively. x htl y htl and z htl These are the x, y, and z coordinates of the trailing edge point of the tail fin tip; Therefore, the key point coordinates of the tail plane are determined by using 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 trailing edge point of the wingtip, the trailing edge point of the wing root, and the fuselage tip point; the key points of the tail plane include the leading edge point of the tail root, the trailing edge point of the tail root, and the trailing edge point of the tailtip. In step S3, using the second optimized aircraft configuration as the initial aircraft shape, the fuselage is modified using the following method to obtain the final low-detonation configuration of the aircraft: Step S3.1: At the Mach number Ma at the aircraft design point, use computational fluid dynamics (CFD) to calculate the constant lift of the current aircraft shape, and extract the near-field overpressure distribution at a distance R directly below the aircraft under the design lift coefficient. And the following formula is used for near-field overpressure distribution. Converted to equivalent cross-sectional area distribution : ； Wherein: γ is the specific heat ratio of 1.2; dp is the static pressure of the infinite inflow; dp is the difference between the local static pressure and the static pressure of the infinite inflow. For the aircraft along the longitudinal axis of the fuselage Length; As variables, The value range is 0~ ; Step S3.2: Obtain the near-field overpressure target distribution with the minimum ground perceived noise level (PLdB) using an optimization method. Step S3.2.1, the near-field overpressure distribution obtained in step S3.1 Parameterization was performed, and the near-field overpressure distribution was selected. The maximum and minimum values on the signal waveform are used as key control points 1 and 2, respectively, and parameterized as piecewise linear functions; where, the key control points The coordinates are ( , ), , and These represent the key control points. The horizontal and vertical coordinates; the horizontal coordinate is along the longitudinal axis of the aircraft, and the vertical coordinate is in the vertical direction; Step S3.2.2, convert the equivalent cross-sectional area distribution in step S3.
1. terminal value As an indicator of aircraft lift, the terminal value is constrained in the optimization. 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 positions 1, 2, and 3; at each key position Place, Its converted equivalent cross-sectional area distribution The following two equations must be satisfied: ； ； in: and They are at key positions. The upper and lower limits of the equivalent cross-sectional area distribution at the point of conversion; For the current optimized configuration at key positions The volume equivalent cross-sectional area distribution at the location; and They are at key positions. The upper and lower limits of the volume equivalent cross-sectional area distribution at the location; Key control points in steps S3.2.4 and S3.2.1 x-coordinate Fixed, with the vertical axis As a design variable; near-field overpressure distribution The generalized Burgers equation is propagated to the ground, and the perceived ground noise level (PLdB) is calculated. Using the perceived ground noise level (PLdB) as the objective function, the ordinate is optimized under the constraints of steps S3.2.2 and S3.2.
3. To minimize the ground-perceived noise level PLdB, the equivalent cross-sectional area distribution corresponding to the optimized solution is the near-field overpressure target distribution. Step S3.2.5: Convert the near-field overpressure target distribution obtained from the optimization in step S3.2.4 into a converted equivalent cross-sectional area target distribution. ; Step S3.3: Adjust the fuselage shape so that the converted equivalent cross-sectional area distribution at point R directly below the fuselage matches the target converted equivalent cross-sectional area distribution. Matching is performed to obtain the aircraft shape after inner-layer iterative design; Step S3.3.1, based on the transformed equivalent cross-sectional area distribution of the current aircraft shape. Volume equivalent cross-sectional area distribution and conversion of equivalent cross-sectional area target distribution The target distribution of volume equivalent cross-sectional area is obtained according to the following mixed confidence approximation formula. ; ； Step S3.3.2: Parameterize the current aircraft shape and analyze the volumetric equivalent cross-sectional area distribution of the aircraft shape. And calculate the target distribution with equivalent cross-sectional area to the volume. The difference of squares is used to obtain the difference of squares of the equivalent cross-sectional area of the volume; Step S3.3.3: Using the minimization of the difference in the square of the equivalent cross-sectional area of the volume as the objective function, optimize the parameters of the aircraft shape until the difference in the square of the equivalent cross-sectional area of the volume cannot be further reduced, and obtain the 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 at position R directly below the aircraft. And converted into the equivalent cross-sectional area distribution. The equivalent cross-sectional area distribution of the transformation was calculated. and conversion of equivalent cross-sectional area target distribution The squared difference is used to obtain the squared difference of the equivalent cross-sectional area. Using the minimization of the squared difference of the equivalent cross-sectional area as the objective function, the parameters of the aircraft shape are optimized until the squared difference of the equivalent cross-sectional area cannot be further reduced, thus obtaining the optimized aircraft shape, which is the aircraft shape after the inner iterative design. Step S3.4: Determine whether the lift coefficient of the aircraft shape after the inner iterative design has converged to the required lift coefficient. If not, return to step S3.1 and repeat steps S3.1 to S3.3 until the lift coefficient of the aircraft shape after the inner iterative design converges to the required lift coefficient. The shape obtained at this point is the final low-sonic-boom configuration of the aircraft designed using the integrated design method of the supersonic civil aircraft low-sonic-boom concept layout.