CN105589995A - Aerodynamic layout optimization method for low sonic boom of supersonic aircraft - Google Patents

Abstract

The invention provides an aerodynamic layout optimization method for a low sonic boom of a supersonic aircraft. The method comprises the following steps: (1) obtaining a geometric layout information table of various components for constructing the aircraft, and calculating a geometric parameter and an aerodynamic parameter value of the aircraft corresponding to each geometric layout information; (2) selecting the geometric layout information of which the geometric parameter conforms to a geometric parameter threshold range and the aerodynamic parameter value conforms to an aerodynamic parameter threshold range as candidate layout information; and (3) calculating A-weighted sound level values of ground sonic booms of the corresponding aircraft in different flight conditions according to each candidate layout information. The geometric layout information corresponding to the minimum A-weighted sound level value is an aerodynamic layout scheme for the low sonic boom of the aircraft. According to the method, with the geometric layout parameter of the aircraft as an optimal design variable, the aerodynamic layout, meeting the aerodynamic characteristics and the noise level requirements, for the low sonic boom of the supersonic aircraft is obtained.

Description

A kind of moving layout optimization method of supersonic aircraft bass gas explosion

Technical field

The invention belongs to aircraft field, particularly a kind of moving layout optimization method of supersonic aircraft bass gas explosion.

Background technology

High sonic boom level is that restriction SST supersonic transport aircraft extensively drops into one of bottleneck of Civil Aviation Market. OneThe sonic boom that 16000 meters of high-altitudes produce taking the Concorde of twice velocity of sound flight is as 133 decibels, and active service seating plane takes off and marches into the arenaNoise only has 90 decibels of left and right. High sonic boom level has directly caused Concorde aircraft to be prohibited across continent supersonic flight, this utmost pointEarth effect the economy of supersonic speed passenger plane. The noise requirements that industrial quarters proposes for SST supersonic transport aircraft of future generation is 70 pointsBelow shellfish, close with the noise level of conventional airplane. Therefore reducing sonic boom level is SST supersonic transport aircraft development process of future generationIn one of key issue urgently to be resolved hurrily.

Generally with the sonic boom based on classical supersonic speed linearized theory to the low sonic boom layout designs of supersonic aircraft at presentMinimizing theoretical is design criteria. This theory is a kind of anti-method for designing, distributes to obtain aircraft by the overvoltage of design sonic boomEquivalent area distributes. The sonic boom problem of supersonic aircraft is one and relates to aircraft layout designs, aerodynamics and acoustics etc.The complicated research field of multiple subjects, adopts sonic boom to minimize the theoretical aircraft layout obtaining and can only meet sonic boom optimum, noCan consider the aspect requirements such as aerodynamic characteristic, aircraft total arrangement, structural design simultaneously. In addition, this theory can only obtain and set upEnumeration sonic boom optimal location, is difficult to meet the designing requirement of many design points, therefore in aircraft layout designs, is had being difficult toEffect is used.

Summary of the invention

An object of the present invention is to solve at least the problems referred to above or defect, and provide at least below by the advantage of explanation.

A further object of the invention is to provide a kind of moving layout optimization method of supersonic aircraft bass gas explosion, to flyRow device geometric layout parameter is optimal design variable, taking aerodynamic characteristics of vehicle as optimizing constraint, by multiple state of flight weightingsSonic boom level, as optimal design target, utilizes modern optimization algorithm to carry out multidisciplinary multiple-objection optimization to aircraft layout, obtainsMust meet the quick-fried aerodynamic arrangement of bass of aerodynamic characteristic and noise level requirement.

In order to realize according to these objects of the present invention and other advantage, provide a kind of supersonic aircraft bass gas explosionMoving layout optimization method, comprising:

A kind of moving layout optimization method of supersonic aircraft bass gas explosion, is characterized in that, comprising:

Step 1, obtain the geometric layout information table of all parts that builds aircraft, and calculate each described geometry and divideThe geometric parameter of the corresponding aircraft of cloth information and pneumatic parameter value

Step 2, default described in geometric parameter threshold range and the aerodynamic parameter threshold range of aircraft to be built; ChoosingGet that described geometric parameter meets described geometric parameter threshold range and described aerodynamic parameter value meets described aerodynamic parameter threshold valueThe geometric layout information of scope is as candidate's layout information;

Step 3, calculate aircraft corresponding thereto at the difference bar that flies according to each described candidate's layout informationThe A weighting weighted sound level of the ground sonic boom under part;

Wherein, the minimum corresponding described geometric layout information of described A weighting weighted sound level is the low of aircraftSonic boom aerodynamic arrangement scheme. From geometric layout information table, choose how much distributed intelligences that meet aerodynamic characteristics of vehicle, then profitCalculate adding of the aircraft corresponding thereto ground sonic boom under different flying conditions with how much distributed intelligences that selectPower A-weighted sound level, chooses the constructing plan of corresponding how much distributed intelligences of minimum A weighting weighted sound level as aircraft,The aircraft that utilizes this constructing plan to build reaches and meets aerodynamic characteristics of vehicle and low noise requirement simultaneously.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, described geometric layout informationRefer to the geometry information of each described parts and described in each parts at carry-on relative position information. Utilize how muchEach described geometric layout information in layout information table can obtain the formalness layout of whole aircraft.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, described step 3 also comprises,Under each flying condition, calculate the A-weighted sound level of the corresponding aircraft of each described candidate's layout information; Giving each fliesRow condition one weight coefficient, obtains the A weighting weighted sound level of this aircraft.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, described A-weighted sound levelComputational process is:

Calculate in aircraft airflight process the near field taking n times of height of aircraft in the spatial dimension of radiusPressure:

$\mathrm{\δ}p(x-\mathrm{\β}r,r)=p_0\frac{{\mathrm{\γM}}^{2}F(x-\mathrm{\β}r)}{{\left(2\mathrm{\β}\mathrm{\γ}\right)}^{1/2}}$

$F\left(y\right)=\frac{1}{2\mathrm{\π}}{\∫}_0^y\frac{{A_e}^{\′\′}\left(\mathrm{\ξ}\right)}{{(y-\mathrm{\ξ})}^{1/2}}d\mathrm{\ξ}$

Wherein: taking aircraft fuselage direction as X-axis, tail direction is forward, is Y-axis, with fuselage perpendicular to fuselage directionTop is forward structure reference axis; δ_pFor certain point pressure p in described spatial dimension is with respect to infinite point pressure p₀Overpressure value; y＝x-βr，M is flight Mach number; γ is specific heat ratio, and x, y are this some x axial coordinate value and y axle in reference axisCoordinate figure, r is the distance value of this point and described aircraft; A_eFor equivalent area;

By near field pressure is carried out to partial differential calculating, obtain the pressure signal that is applied to ground:

$\frac{{\mathrm{dm}}_i}{dt}=\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i^2+\frac{1}{2}(\frac{3}{a_0}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt})m_i;$

$\frac{{\mathrm{d\Δp}}_i}{dt}=\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}(m_i-m_{i-1}){\mathrm{\Δp}}_i+\frac{1}{2}(\frac{3}{a_0}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt}){\mathrm{\Δp}}_i;$

$\frac{{\mathrm{d\λ}}_i}{dt}=-\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}({\mathrm{\Δp}}_i-{\mathrm{\Δp}}_{i+1})-\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i{\mathrm{\λ}}_i$

$m_i=\∂p/\∂T$

λ_i＝T_i+l-T_i；

Wherein, be discrete distribution by geostatic pressure, i and i+1 represent respectively the discrete distributed points of geostatic pressure; m_iRepresent pressureSignal waveform slope; λ_iRepresent the time that each pressure signal is lasting; T_iAnd T_i+1The pressure signal that represents ground is applied to groundTime; △ p_iRepresent pressure signal enhancing amount; ρ₀And a₀Represent respectively atmospheric density value and acoustic velocity value; c_nRepresent that pressure is along rippleThe normal direction spread speed of front; S is the wave surface area that four adjacent pressure rays surround.

Geostatic pressure signal is carried out to Fourier transformation and obtains arrowband pressure distribution:

$P\left(k\right)=\frac{I}{N}{\Σ}_{n=0}^{N-1}p\left(n\right)e^{-jkn\frac{2\mathrm{\π}}{N}},k=0,1,2,...,N-1$

Wherein, p (n) is the pressure signal value on a certain moment ground; Corresponding force value when p (k) is K for frequency; N isData point number, gets 2 exponential.

Calculate the sound pressure level under each frequency:

$SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i,H_i}\left(x_k\right)=20{\log }_{10}\frac{p_e}{p_{ref}}$

SPL_A is the A-weighted sound level of sonic boom noise, Ma_i,Cl_i,H_iRepresent the flight status parameter of aircraft, Ma is MachNumber, Cl is lift coefficient, H is flying height; p_eFor effective acoustic pressure, p_refFor reference sound pressure, get 2 × 10^-5Pa.；

According to 1/3 octave band filtering table, arrowband sound pressure level is converted to third-octave sound pressure level again; Ring according to A weighted againShould calculate A-weighted sound level with the relation of frequency.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, when described aircraft is axleWhen symmetrical structure, A_eFor Mach cone and fuselage method of section are to projected area;

When described aircraft is asymmetric body, A_eComprise two parts: Mach cone and fuselage method of section are to projected area and literForce component A_L(x，θ)；

A_LThe computing formula of (x, θ) is:

$A_L(x,\mathrm{\θ})=\frac{\mathrm{\β}}{{\mathrm{\ρu}}_{\mathrm{\∞}}^2}{\∫}_0^{x}L(x,\mathrm{\θ})dx$

L (x, θ) is the lift component of θ place, X axis position unit length.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, in described step 3, calculateEach state of flight weight coefficient calculates A weighting weighted sound level, is specially:

${\Σ}_{i=1}^N{\mathrm{\ω}}_{i}SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i{H}_i}\left(x_k\right);$

Wherein, ω_iFor the weight coefficient of each design point.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, described in described step 2When aerodynamic characteristics meets aerodynamic characteristics constraints, for:

F_aerodynamic(x_k)∈R_aerodynamic，x_k∈Ω；

Wherein, x_kFormalness data, k is the number of formalness data class; F_aerodynamic(x_k) be formalnessData are x_kTime aircraft aerodynamic characteristic; R_aerodynamicFor aerodynamic characteristics constraints.

Preferably, in the moving layout optimization method of described supersonic aircraft bass gas explosion, described aerodynamic parameter thresholdValue scope comprises lift-drag ratio threshold range, stall angle threshold range and torque factor threshold range.

Beneficial effect of the present invention is as follows:

1, in the moving layout optimization method of described supersonic aircraft bass gas explosion, simultaneously to supersonic aircraft aerodynamic forceCharacteristic and sonic boom level are optimized, and obtain the aerodynamic arrangement that meets aerodynamic characteristic and noise level requirement.

2, in the moving layout optimization method of described supersonic aircraft bass gas explosion, how much to multiple parts simultaneouslyLayout information is optimized, and realizes fast the low sonic boom optimization to multiple design points.

Brief description of the drawings

Fig. 1 is the flow chart of the moving layout optimization method of supersonic aircraft bass gas explosion of the present invention;

Fig. 2 is super in the moving layout optimization method of the supersonic aircraft bass gas explosion described in one of them embodiment of the present inventionVelocity of sound linearized theory schematic diagram.

Detailed description of the invention

Below in conjunction with accompanying drawing, the present invention is described in further detail, to make those skilled in the art with reference to description literary compositionWord can be implemented according to this.

The invention discloses a kind of moving layout optimization method of supersonic aircraft bass gas explosion, as depicted in figs. 1 and 2, shouldMethod at least comprises:

Step 1, obtain the geometric layout information table of all parts that builds aircraft, and calculate each described geometry and divideThe geometric parameter of the corresponding aircraft of cloth information and pneumatic parameter value; Described geometric layout information refers to each described partsDescribed in geometry information and each, parts are at carry-on relative position information.

Step 2, default described in geometric parameter threshold range and the aerodynamic parameter threshold range of aircraft to be built; ChoosingGet that described geometric parameter meets described geometric parameter threshold range and described aerodynamic parameter value meets described aerodynamic parameter threshold valueThe geometric layout information of scope is as candidate's layout information;

Step 3,

3.1), under each flying condition, calculate the A weighted sound of the corresponding aircraft of each described candidate's layout informationLevel;

Calculate in aircraft airflight process the near field taking n times of height of aircraft in the spatial dimension of radiusPressure:

$\mathrm{\δ}p(x-\mathrm{\β}r,r)=p_0\frac{{\mathrm{\γM}}^{2}F(x-\mathrm{\β}r)}{{\left(2\mathrm{\β}\mathrm{\γ}\right)}^{1/2}}$

$F\left(y\right)=\frac{1}{2\mathrm{\π}}{\∫}_0^y\frac{{A_e}^{\′\′}\left(\mathrm{\ξ}\right)}{{(y-\mathrm{\ξ})}^{1/2}}d\mathrm{\ξ}$

Wherein: taking aircraft fuselage direction as X-axis, tail direction is forward, is Y-axis, with fuselage perpendicular to fuselage directionTop is forward structure reference axis; δ_pFor certain point pressure p in described spatial dimension is with respect to infinite point pressure p₀Overpressure value; y＝x-βr，M is flight Mach number; γ is specific heat ratio, and x, y are this some x axial coordinate value and y axle in reference axisCoordinate figure, r is the distance value of this point and described aircraft; A_eFor equivalent area;

By near field pressure is carried out to partial differential calculating, obtain the pressure signal that is applied to ground:

$\frac{{\mathrm{dm}}_i}{dt}=\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i^2+\frac{1}{2}(\frac{3}{a_0}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt})m_i;$

$\frac{{\mathrm{d\Δp}}_i}{dt}=\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}(m_i-m_{i-1}){\mathrm{\Δp}}_i+\frac{1}{2}(\frac{3}{a_0}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt}){\mathrm{\Δp}}_i;$

$\frac{{\mathrm{d\λ}}_i}{dt}=-\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}({\mathrm{\Δp}}_i-{\mathrm{\Δp}}_{i+1})-\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i{\mathrm{\λ}}_i$

$m_i=\∂p/\∂T$

λ_i＝T_i+l-T_i；

Wherein, be discrete distribution by geostatic pressure, i and i+1 represent respectively the discrete distributed points of geostatic pressure; m_iRepresent pressureSignal waveform slope; λ_iRepresent the time that each pressure signal is lasting; T_iAnd T_i+1The pressure signal that represents ground is applied to groundTime; △ p_iRepresent pressure signal enhancing amount; ρ₀And a₀Represent respectively atmospheric density value and acoustic velocity value; c_nRepresent that pressure is along rippleThe normal direction spread speed of front; S is the wave surface area that four adjacent pressure rays surround.

Geostatic pressure signal is carried out to Fourier transformation and obtains arrowband pressure distribution:

$P\left(k\right)=\frac{I}{N}{\Σ}_{n=0}^{N-1}p\left(n\right)e^{-jkn\frac{2\mathrm{\π}}{N}},k=0,1,2,...,N-1$

Wherein, p (n) is the pressure signal value on a certain moment ground; Corresponding force value when p (k) is K for frequency; N isData point number, gets 2 exponential.

Calculate the sound pressure level under each frequency:

$SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i,H_i}\left(x_k\right)=20{\log }_{10}\frac{p_e}{p_{ref}}$

SPL_A is the A-weighted sound level of sonic boom noise, Ma_i，Cl_i，H_iRepresent the flight status parameter of aircraft, Ma is MachNumber, Cl is lift coefficient, H is flying height; p_eFor effective acoustic pressure, p_refFor reference sound pressure, get 2 × 10^-5Pa.；

According to 1/3 octave band filtering table as shown in table 1, arrowband sound pressure level is converted to third-octave sound pressure level again;

The filtering of table 11/3 octave band

1/3 octave band centre frequency/Hz	Lower limit	The upper limit	1/3 octave band centre frequency/Hz	Lower limit	The upper limit
						25	22.4	28	800	710	900
31.5	28	35.5	1000	900	1120
						40	35.5	45	1250	1120	1400
50	45	56	1600	1400	1800
						63	56	71	2000	1800	2240
80	71	90	2500	2240	2800
						100	90	112	3150	2800	3550
125	112	144	4000	3550	4500
						160	144	180	5000	4500	5600
200	180	224	6300	5600	7100
						250	224	280	8000	7100	9000
315	280	355	10000	9000	11200
						400	355	450	12500	11200	14000
500	450	560	16000	14000	18000
						630	560	710	20000	18000	22400

Calculate A-weighted sound level according to A weighted response as shown in table 2 with the relation of frequency again.

The relation of table 2A weighted response and frequency

Frequency/Hz	A weighted correction/dB	Frequency/Hz	A weighted correction/dB
				20	-50.5	630	-1.9
25	-44.7	800	-0.8
				31.5	-39.4	1000	0
40	-34.6	1250	0.6
				50	-30.2	1600	1.0 5 -->
63	-26.2	2000	1.2
				80	-22.5	2500	1.3
100	-19.1	3150	1.2
				125	-16.1	4000	1.0
160	-13.4	5000	0.5
				200	-10.9	6300	-0.1
250	-8.6	8000	-1.1
				315	-6.6	10000	-2.5
400	-4.8	12500	-4.3
				500	-3.2	16000	-6.6

3.2) give each flying condition one weight coefficient, obtain the A weighting weighted sound level of this aircraft.

${\Σ}_{i=1}^N{\mathrm{\ω}}_{i}SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i{H}_i}\left(x_k\right);$

Wherein, ω_iFor the weight coefficient of each design point.

Wherein, the minimum corresponding described geometric layout information of described A weighting weighted sound level is the low of aircraftSonic boom aerodynamic arrangement scheme.

In such scheme, in the time that described aircraft is axially symmetric structure, A_eFor Mach cone and fuselage method of section are to perspective planeLong-pending;

When described aircraft is asymmetric body, A_eComprise two parts: Mach cone and fuselage method of section are to projected area and literForce component A_L(x，θ)；

A_LThe computing formula of (x, θ) is:

$A_L(x,\mathrm{\θ})=\frac{\mathrm{\β}}{{\mathrm{\ρu}}_{\mathrm{\∞}}^2}{\∫}_0^{x}L(x,\mathrm{\θ})dx$

L (x, θ) is the lift component of θ place, X axis position unit length.

In such scheme, when described in described step 2, aerodynamic characteristics meets aerodynamic characteristics constraints, for:

F_aerodynamic(x_k)∈R_aerodynamic，x_k∈Ω；

Wherein, x_kFormalness data, k is the number of formalness data class; F_aerodynamic(x_k) be formalnessData are x_kTime aircraft aerodynamic parameter value; R_aerodynamicFor aerodynamic parameter threshold range.

In such scheme, described aerodynamic parameter threshold range comprises lift-drag ratio threshold range, stall angle threshold rangeWith torque factor threshold range.

Although embodiments of the invention are open as above, it is not restricted to listed fortune in description and embodimentWith, it can be applied to various applicable the field of the invention completely, for those skilled in the art, and can be easily realNow other amendment, is not therefore deviating under the universal that claim and equivalency range limit, and the present invention is not limited toSpecific details.

Claims (8)

1. the moving layout optimization method of supersonic aircraft bass gas explosion, is characterized in that, comprising:

Step 1, obtain the geometric layout information table of all parts that builds aircraft, and calculate each described letter that distributes for how muchCease geometric parameter and the pneumatic parameter value of corresponding aircraft;

Step 2, default described in geometric parameter threshold range and the aerodynamic parameter threshold range of aircraft to be built; Choose instituteState that geometric parameter meets described geometric parameter threshold range and described aerodynamic parameter value meets described aerodynamic parameter threshold rangeGeometric layout information as candidate's layout information;

Step 3, calculate aircraft corresponding thereto under different flying conditions according to each described candidate's layout informationThe A weighting weighted sound level of ground sonic boom;

Wherein, the minimum corresponding described geometric layout information of described A weighting weighted sound level is the low sonic boom of aircraftAerodynamic arrangement's scheme.

2. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 1, is characterized in that described how muchLayout information refer to the geometry information of each described parts and described in each parts at carry-on relative position information.

3. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 2, is characterized in that described stepThree also comprise, under each flying condition, calculate the A-weighted sound level of the corresponding aircraft of each described candidate's layout information;Give each flying condition one weight coefficient, obtain the A weighting weighted sound level of this aircraft.

4. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 3, is characterized in that, described A meterThe computational process of power sound level is:

Calculate in aircraft airflight process the near field pressure taking n times of height of aircraft in the spatial dimension of radius:

$\begin{array}{c}\mathrm{\δ}p\left(x-\mathrm{\β}r,r\right)=p_0\frac{{\mathrm{\γM}}^{2}F\left(x-\mathrm{\β}r\right)}{{\left(2\mathrm{\β}\mathrm{\γ}\right)}^{1/2}}\\ F\left(y\right)=\frac{1}{2\mathrm{\π}}{\∫}_0^y\frac{{\mathrm{Ae}}^{\′\′}\left(\mathrm{\ξ}\right)}{{\left(y-\mathrm{\ξ}\right)}^{1/2}}d\mathrm{\ξ}\end{array}$

Wherein: taking aircraft fuselage direction as X-axis, tail direction is forward, is Y-axis, with body upper perpendicular to fuselage directionFor forward builds reference axis; δ_pFor certain point pressure p in described spatial dimension is with respect to infinite point pressure p₀Overpressure value; Y=x-βr，M is flight Mach number; γ is specific heat ratio, and x, y are this some x axial coordinate value and y axial coordinate in reference axisValue, r is the distance value of this point and described aircraft; A_eFor equivalent area;

By near field pressure is carried out to partial differential calculating, obtain the pressure signal that is applied to ground:

$\frac{{\mathrm{dm}}_i}{dt}=\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i^2+\frac{1}{2}(\frac{3}{a_0}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt})m_i;$

$\frac{{\mathrm{d\Δp}}_i}{dt}=\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}(m_i-m_{i-1}){\mathrm{\Δp}}_i+\frac{1}{2}(\frac{3}{a_o}\frac{{\mathrm{da}}_0}{dt}+\frac{1}{{\mathrm{\ρ}}_0}\frac{{\mathrm{d\ρ}}_0}{dt}-\frac{2}{c_n}\frac{{\mathrm{dc}}_n}{dt}-\frac{1}{S}\frac{dS}{dt}){\mathrm{\Δp}}_i;$

$\frac{{\mathrm{d\λ}}_i}{dt}=-\frac{\mathrm{\γ}+1}{4\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}({\mathrm{\Δp}}_i-{\mathrm{\Δp}}_{i+1})-\frac{\mathrm{\γ}+1}{2\mathrm{\γ}}\frac{a_0}{{\mathrm{\ρ}}_0{c}_n}{m}_i{\mathrm{\λ}}_i$

$m_i=\∂p/\∂T$

λ_i＝T_i+1-T_i；

Wherein, be discrete distribution by geostatic pressure, i and i+1 represent respectively the discrete distributed points of geostatic pressure; m_iRepresent pressure signalWaveform slope; λ_iRepresent the time that each pressure signal is lasting; T_iAnd T_i+1Represent the pressure signal on ground be applied to ground timeBetween; Δ p_iRepresent pressure signal enhancing amount; ρ₀And a₀Represent respectively atmospheric density value and acoustic velocity value; Cn represents that pressure is along wave surfaceNormal direction spread speed; S is the wave surface area that four adjacent pressure rays surround.

Geostatic pressure signal is carried out to Fourier transformation and obtains arrowband pressure distribution:

$P\left(k\right)=\frac{1}{N}{\Σ}_{n=0}^{N-1}p\left(n\right)e^{-jkn\frac{2\mathrm{\π}}{N}},k=0,1,2,...,N-1$

Wherein, p (n) is the pressure signal value on a certain moment ground; Corresponding force value when p (k) is K for frequency; N is dataPoint number, gets 2 exponential.

Calculate the sound pressure level under each frequency:

$SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i,H_i}\left(x_k\right)=20{\log }_{10}\frac{p_e}{p_{ref}}$

SPL_A is the A-weighted sound level of sonic boom noise, Ma_i,Cl_i,H_iThe flight status parameter that represents aircraft, Ma is Mach number,Cl is lift coefficient, and H is flying height; p_eFor effective acoustic pressure, p_refFor reference sound pressure, get 2 × 10^-5Pa.；

According to 1/3 octave band filtering table, arrowband sound pressure level is converted to third-octave sound pressure level again; Again according to A weighted response withThe relation of frequency calculates A-weighted sound level.

5. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 4, is characterized in that,

In the time that described aircraft is axially symmetric structure, A_eFor Mach cone and fuselage method of section are to projected area;

When described aircraft is asymmetric body, A_eComprise two parts: Mach cone and fuselage method of section are to projected area and lift componentA_L(x，θ)；

A_LThe computing formula of (x, θ) is:

$A_L(x,\mathrm{\θ})=\frac{\mathrm{\β}}{{\mathrm{\ρu}}_{\mathrm{\∞}}^2}{\∫}_0^{x}L\left(x,\mathrm{\θ}\right)dx$

L (x, θ) is the lift component of θ place, X axis position unit length.

6. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 5, is characterized in that described stepIn three, calculate each state of flight weight coefficient and calculate A weighting weighted sound level, be specially:

${\Σ}_{i=1}^N{\mathrm{\ω}}_{i}SPL\_A_{{\mathrm{Ma}}_i,{\mathrm{Cl}}_i,H_i}\left(x_k\right);$

Wherein, ω_iFor the weight coefficient of each design point.

7. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 6, is characterized in that described stepWhen aerodynamic characteristics described in two meets aerodynamic characteristics constraints, for:

F_aerodynamic(x_k)∈R_aerodynamic，x_k∈Ω；

Wherein, x_kFormalness data, k is the number of formalness data class; F_aerodynamic(x_k) for formalness data bex_kTime aircraft aerodynamic characteristic; R_aerodynamicFor aerodynamic characteristics constraints.

8. the moving layout optimization method of supersonic aircraft bass gas explosion as claimed in claim 6, is characterized in that, described pneumaticForce parameter threshold range comprises lift-drag ratio threshold range, stall angle threshold range and torque factor threshold range.