Wide-speed-range atmospheric parameter resolving method and electronic equipment

Abstract

The embodiment of the invention relates to the field of waverider configuration wide-area aircrafts, and discloses a wide-speed-range atmospheric parameter resolving method and electronic equipment.The wide-speed-range atmospheric parameter resolving method comprises the steps that pressure data and a reference Mach number of a preset pressure measuring point on a target waverider configuration wide-area aircraft are obtained; determining a target resolving mode of atmospheric parameters by using the reference Mach number, the target resolving mode including executing one or two algorithms of a subsonic velocity algorithm, a transsonic velocity algorithm and a hypersonic velocity algorithm to resolve the atmospheric parameters, and the atmospheric parameters including an attack angle, a sideslip angle and the Mach number; and resolving atmospheric parameters in the target resolving mode. According to the method, the problem of resolving the full-speed-domain atmospheric parameters of the wave-rider-configuration wide-domain aircraft is solved, and in the resolving process, resolving modes can be rapidly switched according to actual conditions, so that accurate resolving of flight parameters such as the attack angle, the sideslip angle and the Mach number in the whole flight process of the wave-rider-configuration wide-domain aircraft is achieved.

Description

Technical Field

[0001] The embodiments of the present invention relate to the field of waverider configuration wide - domain aircraft, and particularly to a wide - speed - domain atmospheric parameter calculation method and an electronic device. Background Art

[0002] As a frontier exploration in the fields of aerospace and weaponry at present, the waverider wide - domain aircraft is a kind of aircraft with characteristics such as wide speed domain, high altitude, and strong penetration ability. To achieve efficient and stable flight of the aircraft, accurate flight parameter inputs are required to achieve precise and real - time control of the wide - domain aircraft.

[0003] Traditional low - speed aircraft usually use pitot tubes or embedded atmospheric data sensing systems to measure and calculate flight parameters. However, in the extreme environment of high Mach - number flight, the probe - type sensors protruding from the aircraft surface are prone to failure, and inertial navigation systems or embedded atmospheric data sensing systems are usually used to obtain flight parameters such as airspeed, angle of attack, and sideslip angle.

[0004] At present, the embedded atmospheric data sensing system has been widely studied and applied in engineering. For example, aircraft such as B - 2, F18, and X - 38 are all applicable. However, current research usually focuses on the measurement and calculation of atmospheric parameters in the low - speed to supersonic speed range. On the one hand, the embedded atmospheric data sensing system is greatly affected by airflow interference, especially under conditions of large angle of attack and high dynamic flight, and its measurement accuracy decreases. On the other hand, the flight speed range of the waverider configuration wide - domain aircraft covers from low speed to hypersonic speed. Due to the different effects of air compression performance and strong shock waves in different speed ranges, it is difficult to ensure the calculation stability and adaptability of the embedded atmospheric data calculation model in the full speed range. Summary of the Invention

[0005] The purpose of the present invention is to provide at least a wide - speed - domain atmospheric parameter calculation method and an electronic device, which can at least solve the problem of calculating atmospheric parameters in the full speed range of the waverider configuration wide - domain aircraft. During the calculation process, the calculation mode can be quickly switched according to the actual situation to achieve accurate calculation of flight parameters such as angle of attack, sideslip angle, and Mach number throughout the flight process of the waverider configuration wide - domain aircraft.

[0006] The embodiments of the present invention provide a wide - speed - domain atmospheric parameter calculation method, including:

[0007] Obtain the pressure data and reference Mach number of preset pressure measurement points on the target waverider configuration wide - domain aircraft;

[0008] Determine the target calculation mode of atmospheric parameters by using the reference Mach number. The target calculation mode includes performing one or two of the subsonic algorithm, trans - supersonic algorithm, and hypersonic algorithm to calculate atmospheric parameters. The atmospheric parameters include angle of attack, sideslip angle, and Mach number;

[0009] Calculate the atmospheric parameters in the target calculation mode.

[0010] In some alternative embodiments, the preset pressure measurement points are the pressure measurement points where the pressure regularly changes on the upper and lower surfaces of the target waverider configuration wide-body aircraft.

[0011] In some alternative embodiments, determining the target calculation mode of the atmospheric parameters using the reference Mach number includes:

[0012] When the reference Mach number is within the first Mach number range, execute the subsonic algorithm; when the reference Mach number is within the second Mach number range, execute the subsonic algorithm and the transonic algorithm simultaneously; when the reference Mach number is within the third Mach number range, execute the transonic algorithm; when the reference Mach number is within the fourth Mach number range, execute the transonic algorithm and the hypersonic algorithm simultaneously; when the reference Mach number is greater than the upper limit of the fourth Mach number range, execute the hypersonic algorithm;

[0013] Wherein, the second Mach number range exceeds the upper limit of the first Mach number range, the third Mach number range exceeds the upper limit of the second Mach number range, and the fourth Mach number range exceeds the upper limit of the third Mach number range.

[0014] In some alternative embodiments, the preset pressure measurement points include the first pressure measurement point, the second pressure measurement point, the third pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the eighth pressure measurement point located on the upper surface of the head of the target waverider configuration wide-body aircraft, and the ninth pressure measurement point and the tenth pressure measurement point located on the lower surface of the head of the fuselage of the target waverider configuration wide-body aircraft;

[0015] Combine the pressure data of the third pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the tenth pressure measurement point for calculating the atmospheric parameters by the subsonic algorithm, combine the pressure data of the first pressure measurement point, the second pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the ninth pressure measurement point for calculating the atmospheric parameters by the transonic algorithm, and combine the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point for calculating the atmospheric parameters by the hypersonic algorithm.

[0016] In some alternative embodiments, the subsonic algorithm includes:

[0017] Extract the pressure data of the third pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the tenth pressure measurement point;

[0018] Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP3 at the third pressure measurement point, and the pressure coefficient CpsP8 at the eighth pressure measurement point. Cpa0 = (P3 - P8) / (P3 - P10), Cpb0 = (P4 - P5) / (P3 - P8), CpsP3 = (P3 - P s ) / (P t - P s ), CpsP8 = (P8 - P s ) / (P t - P s ), where P3 represents the pressure data at the third pressure measurement point, P8 represents the pressure data at the eighth pressure measurement point, P10 represents the pressure data at the tenth pressure measurement point, P4 represents the pressure data at the fourth pressure measurement point, P5 represents the pressure data at the fifth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0019] Set the initial Mach number within the first Mach number range and start the Mach number iterative solution loop;

[0020] During the Mach number solution loop, iteratively solve the angle of attack and sideslip angle based on the pre-established subsonic basic coefficient table;

[0021] When the solved angle of attack and sideslip angle converge, solve the corresponding total pressure and static pressure according to the subsonic basic coefficient table and the current angle of attack and sideslip angle, and use the solved total pressure and static pressure to calculate the current Mach number;

[0022] When the solved current Mach number converges, output the current angle of attack, sideslip angle, and Mach number.

[0023] In some alternative embodiments, the transonic algorithm includes:

[0024] Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the ninth pressure measurement point;

[0025] Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP1 at the first pressure measurement point, and the pressure coefficient CpsP8 at the eighth pressure measurement point. Cpa0 = (P2 - P9) / (P1 - P8), Cpb0 = (P4 - P5) / (P1 - P8), CpsP1 = (P1 - P s ) / (P t - P s ), CpsP8 = (P8 - P s ) / (P t - P s),P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P8 represents the pressure data of the eighth pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P4 represents the pressure data of the fourth pressure measurement point, P5 represents the pressure data of the fifth pressure measurement point, P t represents the total pressure, P s represents the static pressure;

[0026] Set the initial Mach number within the third Mach number range and start the Mach number iterative solution loop;

[0027] During the Mach number solution loop, iteratively solve the angle of attack and sideslip angle based on a pre-established transonic basic coefficient table;

[0028] When the calculated angle of attack and sideslip angle converge, solve the corresponding total pressure and static pressure according to the transonic basic coefficient table and the current angle of attack and sideslip angle, and use the solved total pressure and static pressure to calculate the current Mach number;

[0029] When the calculated current Mach number converges, output the current angle of attack, sideslip angle, and Mach number.

[0030] In some alternative embodiments, the hypersonic algorithm includes:

[0031] Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point;

[0032] Construct the Mach number characteristic coefficient CpMa0, the angle of attack characteristic coefficient Cpa0, the sideslip angle characteristic coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measurement point, and the pressure coefficient CpsP9 of the ninth pressure measurement point, CpMa0 = P1 / P9, Cpa0 = (P1 - P9) / (P2 - P9), Cpb0 = (P6 - P7) / (P1 - P9), CpsP1 = (P1 - P s ) / (P t -P s ),CpsP9 = (P9 - P s ) / (P t -P s ),P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P6 represents the pressure data of the sixth pressure measurement point, P7 represents the pressure data of the seventh pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P t represents the total pressure, P s represents the static pressure;

[0033] Set the initial barometric altitude and start the barometric altitude iterative solution loop;

[0034] During the air pressure altitude calculation loop, the angle of attack, sideslip angle, and Mach number are iteratively calculated based on a pre-established hypersonic base coefficient table.

[0035] When the calculated angle of attack, sideslip angle, and Mach number converge, the corresponding total pressure and static pressure are solved according to the hypersonic base coefficient table and the current angle of attack, sideslip angle, and Mach number, and the current air pressure altitude is calculated using the solved total pressure and static pressure.

[0036] When the calculated current air pressure altitude converges, the current angle of attack, sideslip angle, and Mach number are output.

[0037] In some alternative embodiments, the subsonic base coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measurement point at a given speed and attitude under subsonic conditions and normalizing the simulated pressure data into dimensionless coefficients; the transonic base coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measurement point at a given speed and attitude under transonic conditions and normalizing the simulated pressure data into dimensionless coefficients; the hypersonic base coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measurement point at a given speed and attitude under hypersonic conditions and normalizing the simulated pressure data into dimensionless coefficients.

[0038] In some alternative embodiments, when performing the air parameter calculation using two of the subsonic algorithm, transonic algorithm, and hypersonic algorithm, after calculating the air parameters in the target calculation mode, it further includes:

[0039] Comparing the Mach numbers in the two sets of air parameters calculated by the two sets of algorithms with a reference Mach number;

[0040] Determining the set of air parameters corresponding to the Mach number with a relatively smaller difference from the reference Mach number as the result of the air parameter calculation.

[0041] An embodiment of the present invention further provides an electronic device, including:

[0042] At least one processor; and,

[0043] A memory communicatively connected to the at least one processor; wherein,

[0044] The memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the above-mentioned wide-speed-domain air parameter calculation method.

[0045] The present invention constructs a full-speed-range atmospheric parameter calculation solution for a waverider-configured wide-range aircraft. During the calculation process, it can quickly switch the calculation mode according to the actual situation to achieve the accurate calculation of flight parameters such as angle of attack, sideslip angle, and Mach number throughout the flight process of the waverider-configured wide-range aircraft, and has good engineering application prospects. Further, based on the atmospheric parameter pressure measurement points of the waverider-configured wide-range aircraft layout, atmospheric parameter calculation algorithms are respectively established for the flow field characteristics in three different speed ranges, namely subsonic, transonic, and hypersonic, to form a wide-range atmospheric data measurement and calculation solution. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] One or more embodiments are exemplarily illustrated by the pictures in the corresponding drawings, and these exemplary illustrations do not limit the embodiments.

[0047] Figure 1 It is a schematic flowchart of the wide-speed-range atmospheric parameter calculation method provided by an embodiment of the present invention;

[0048] Figure 2 It is a schematic diagram of the preset pressure measurement point layout provided by an embodiment of the present invention;

[0049] Figure 3 It is a schematic diagram of the atmospheric parameter calculation principle provided by an embodiment of the present invention;

[0050] Figure 4 It is a schematic diagram of the atmospheric parameter calculation device provided by an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0051] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the embodiments of the present invention will be described in detail below with reference to the drawings. However, those of ordinary skill in the art can understand that in the embodiments of the present invention, many technical details are provided to help readers better understand the present invention. However, even without these technical details and various changes and modifications based on the following embodiments, the technical solutions claimed by the present invention can still be implemented. The following division of each embodiment is for convenience of description and should not constitute any limitation to the specific implementation manner of the present invention. Each embodiment can be combined and cross-referenced with each other without contradiction.

[0052] To facilitate the understanding of the embodiments of the present invention, the atmospheric parameter calculation principle is introduced here first.

[0053] The input of the embedded atmospheric parameter solution algorithm is the pressure measurement value of the pressure measuring point, and the output is the atmospheric parameters such as angle of attack, sideslip angle, Mach number, etc. Usually, the atmospheric parameter solution algorithm is divided into two parts: angle of attack and sideslip angle solution and Mach number solution. First, the measured pressure on the surface is used as the input of the angle of attack and sideslip angle. The angle of attack and sideslip angle can be calculated by calculating the pressure difference relationship between the pressure measuring points; based on the calculated angle of attack and sideslip angle, the static pressure and dynamic pressure are iteratively solved, and finally the Mach number is calculated based on the static pressure and dynamic pressure.

[0054] The following is the process of calculating the angle of attack, sideslip angle, and Mach number.

[0055] Based on the classical solution of the flow around a sphere, the distribution formula of the surface pressure and the equation relationship between the incident angle, the angle of attack and the sideslip angle can be expressed as shown in equations (1) and (2) respectively:

[0056] P i =q c [cos 2 (θ i )+εsin 2 (θ i )]+P s (1)

[0057] cosθ i =cosαcosβcosλ i +sinβsinφ i sinλ i +sinαcosβcosφ i sinλ i (2)

[0058] Where P i is the pressure measurement value of the i-th pressure measuring point; q c is the dynamic pressure; P s is the static pressure, where ε is the shape pressure coefficient, which is related to the angle of attack, sideslip angle, static pressure and dynamic pressure; α is the angle of attack; β is the sideslip angle; φ i is the circumferential angle of the i-th pressure measuring point (defined as the angle between the vertical line of the section where the pressure measuring point is located and the radial direction of the pressure measuring point); i is the cone angle of the ith pressure measuring point (defined as the angle between the normal direction of the pressure measuring point and the horizontal center line); θ i is the incident angle of the ith pressure measuring point (defined as the angle between the surface normal direction of the pressure measuring point and the incoming flow velocity vector).

[0059] Take the pressure values ​​P at three different pressure measuring points i , P j , P k Substituting into equation (1), the simultaneous equations can be simplified to eliminate q c , P s, ε, as shown in Equation (3).

[0060]

[0061] Let the intermediate variable Γ ik = P i - P k , Γ ji = P j - P i , Γ kj = P k - P j , then Equation (4) can be obtained by transforming Equation (3).

[0062] Γ ik cos 2 (θ j ) + Γ ji cos 2 (θ k ) + Γ kj cos 2 (θ i ) = 0 (4)

[0063] Substitute Equation (2) into Equation (4), and divide both sides of the equation by cos 2 β, then Equation (5) can be obtained.

[0064] Γ ik (a j + b j tanβ) 2 + Γ ji (a k + b k tanβ) 2 + Γ kj (a i + b i tanβ) 2 = 0 (5)

[0065] where, a n = cosαcosλ n + sinαcosφ n sinλ n , b n = sinφ n sinλ n (n = i, j, k).

[0066] Since the circumferential angle φ i = 0 (±180°) at the pressure measurement points on the vertical plane of the waverider configuration wide - area aircraft, three points on the vertical line of the waverider configuration wide - area aircraft surface are selected to simplify the angle - of - attack calculation, then Equation (5) is transformed into Equation (6).

[0067] Γ ik (cosαcosλ j +sinαcosφ j sinλ j ) 2

[0068] +Γ ji (cosαcosλ k +sinαcosφ k sinλ k ) 2 (6)

[0069] +Γ kj (cosαcosλ i +sinαcosφ i sinλ i ) 2 =0

[0070] Expand and sort out to get a quadratic equation about tanα (angle of attack solution formula):

[0071] (Γ ik sin 2 λ j +Γ ji sin 2 λ k +Γ kj sin 2 λ i )(tan 2 α - 1)

[0072] +2(Γ ik cosλ j sinλ j cosφ j +Γ ji cosλ k sinλ k cosφ k (7)

[0073] +Γ kj cosλ i sinλ i cosφ i )tanα=0

[0074] Let:

[0075] A = Γ ik sin 2 λ j +Γ ji sin 2 λ k +Γ kj sin 2 λi

[0076] B = Γ ik cosλ j sinλ j cosφ j +Γ ji cosλ k sinλ k cosφ k +Γ kj cosλ i sinλ i cosφ i

[0077] Solve the quadratic equation (6). Considering that the angle of attack is non - negative, we get: when |α| ≤ 45, when |α| > 45, Thus, the angle of attack is calculated.

[0078] Expand and rearrange Equation (5) to get Equation (8).

[0079] (Γ ik b j 2 +Γ ji b k 2 +Γ kj b i 2 )tan 2 β + 2(Γ ik a j b j +Γ ji a k b k +Γ kj a i b i )tanβ(8)

[0080] +Γ ik a j 2 +Γ ji a k 2 +Γ kj a i 2 = 0

[0081] Let:

[0082] A · = Γ ik b j 2 +Γ ji b k 2 +Γkj b i 2 ,

[0083] B · = Γ ik a j b j + Γ ji a k b k + Γ kj a i b i

[0084] C · = Γ ik a j 2 + Γ ji a k 2 + Γ kj a i 2

[0085] Then it can be solved that:

[0086]

[0087] By analyzing the relationship between the sideslip angles β1 and β2, the principle for determining the sideslip angle is as follows:

[0088] (1) When the three pressure measurement points calculated are on the horizontal symmetry line of the waverider configuration wide-body aircraft, the calculated sideslip angle should select the sideslip angle with the smaller median value between the sideslip angles β1 and β2;

[0089] (2) When the three pressure measurement points calculated are not all on the horizontal symmetry line of the waverider configuration wide-body aircraft and the three points are on the same vertical symmetry line, a singular solution situation will occur, and then it is necessary to re-select the pressure measurement points for calculation. In other cases, select the sideslip angle value with the smaller absolute value.

[0090] After completing the above calculations of the angle of attack and the sideslip angle, the dynamic pressure and the static pressure are solved by the iterative method, and the iterative formula is:

[0091]

[0092] where q c is the dynamic pressure; P s is the static pressure; M j , Q are the coefficient matrix and the weight matrix respectively.

[0093]

[0094] After iteratively obtaining the static pressure and dynamic pressure, under subsonic flight conditions, the Mach number Ma can be calculated according to the isentropic flow law in different speed ranges, as shown in Equation (12).

[0095]

[0096] Under supersonic and hypersonic flight conditions, according to the adiabatic normal shock wave relationship, there are:

[0097]

[0098] In addition, for the static pressure P s , dynamic pressure q c and total pressure P t , there is also the following relationship:

[0099] P s +q c =P t (14)

[0100] The barometric altitude can be solved according to the relationship between atmospheric pressure and altitude under hydrostatic equilibrium conditions. The relationship between the two is:

[0101]

[0102] where P s is the static pressure, P d is the ground pressure, R is the specific gas constant of air, T is the temperature, and g is the acceleration due to gravity.

[0103] The present invention constructs a full-speed-range atmospheric parameter calculation scheme for a waverider configuration wide-range aircraft. Based on the atmospheric parameter pressure measurement points of the waverider configuration wide-range aircraft layout, atmospheric parameter calculation algorithms are established respectively for the flow field characteristics in three different speed ranges: subsonic, transonic, and hypersonic. During the calculation process, it can quickly switch according to the actual situation to achieve accurate calculation of flight parameters such as angle of attack, sideslip angle, and Mach number during the entire flight process of the waverider configuration wide-range aircraft, and has good engineering application prospects. The implementation details of the wide-speed-range atmospheric parameter calculation method in this embodiment are specifically described below. The following content is only the implementation details provided for convenient understanding and is not necessary for implementing this solution.

[0104] Embodiment 1

[0105] The wide-speed-range atmospheric parameter calculation method in this embodiment can be applied to an electronic device with communication, computing, and data storage capabilities. Its specific process can be as Figure 1 shown, including:

[0106] Step 101, obtain the pressure data and reference Mach number of the preset pressure measurement points on the target waverider configuration wide-range aircraft.

[0107] In a specific implementation, through CFD (Computational Fluid Dynamics) simulation, the pressure distribution on the surface of a wide - domain aircraft with a target waverider configuration under different flight states can be analyzed, and points with relatively good regularity of surface pressure change are selected as preset pressure - measuring points, that is: the preset pressure - measuring points are the pressure - measuring points with regular pressure change on the upper and lower surfaces of the wide - domain aircraft with a target waverider configuration.

[0108] In one example, the layout of the preset pressure - measuring points is as Figure 2 shown.

[0109] Figure 2 In, the top - view direction of the wide - domain aircraft with a target waverider configuration is shown. The preset pressure - measuring points include the first pressure - measuring point P1, the second pressure - measuring point P2, the third pressure - measuring point P3, the fourth pressure - measuring point P4, the fifth pressure - measuring point P5, the sixth pressure - measuring point P6, the seventh pressure - measuring point P7, and the eighth pressure - measuring point P8 on the upper surface of the head of the wide - domain aircraft with a target waverider configuration, and the ninth pressure - measuring point P9 and the tenth pressure - measuring point P10 on the lower surface of the head of the fuselage of the wide - domain aircraft with a target waverider configuration.

[0110] Among them, the first pressure - measuring point P1, the second pressure - measuring point P2, the third pressure - measuring point P3, the eighth pressure - measuring point P8, and the ninth pressure - measuring point P9 and the tenth pressure - measuring point P10 are located on the intersection line of the vertical symmetry plane of the nose and the surface of the head of the fuselage. The fourth pressure - measuring point P4 and the fifth pressure - measuring point P5 are symmetric about the vertical symmetry plane of the nose, and the sixth pressure - measuring point P6 and the seventh pressure - measuring point P7 are symmetric about the vertical symmetry plane of the nose.

[0111] Specifically, the pressure data of the third pressure - measuring point, the fourth pressure - measuring point, the fifth pressure - measuring point, the eighth pressure - measuring point, and the tenth pressure - measuring point are combined for the solution of atmospheric parameters in the subsonic algorithm:

[0112] 1) P3: Used as the upper - surface pressure for the solution of angle - of - attack pressure difference and as the large - pressure point for the approximate dynamic - pressure difference;

[0113] 2) P4 and P5: Mainly used for the solution of sideslip angle;

[0114] 3) P8, used as the small - pressure point for providing the approximate dynamic - pressure difference;

[0115] 4) P10: Used as the lower - surface pressure for the solution of angle - of - attack pressure difference.

[0116] Specifically, the pressure data of the first pressure - measuring point, the second pressure - measuring point, the fourth pressure - measuring point, the fifth pressure - measuring point, the eighth pressure - measuring point, and the ninth pressure - measuring point are combined for the solution of atmospheric parameters in the trans - supersonic algorithm:

[0117] 1) P1: Used as the large - pressure point for the approximate dynamic - pressure difference;

[0118] 2) P2: Used as the upper surface pressure for angle of attack pressure difference calculation;

[0119] 3) P4 and P5: Mainly used for sideslip angle calculation;

[0120] 4) P8: Used as a small pressure point to provide an approximate dynamic pressure difference;

[0121] 5) P9: Used as the lower surface pressure for angle of attack pressure difference calculation.

[0122] Specifically, the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point are combined for the atmospheric parameter calculation of the hypersonic algorithm:

[0123] 1) P1: Used as a large pressure point for approximate dynamic pressure difference;

[0124] 2) P2: Used as the upper surface pressure for angle of attack pressure difference calculation;

[0125] 3) P4, P5, P6 and P7: Mainly used for sideslip angle calculation;

[0126] 4) P9: Used as the lower surface pressure for angle of attack pressure difference calculation and as a small pressure point for approximate dynamic pressure difference.

[0127] It should be understood that the above pressure measurement points are only examples, and the actual pressure measurement point selection scheme can determine the number and distribution form of the pressure measurement points according to the aerodynamic shape of the aircraft.

[0128] In practical applications, the reference Mach number is determined by means of the output parameters of other devices, such as inertial navigation devices, to perform the solution mode switching so that accurate parameter calculation can be carried out at different flight speeds.

[0129] Step 102, determining the target solution mode of the atmospheric parameters by using the reference Mach number, where the target solution mode includes performing one or two of the subsonic algorithm, the transonic algorithm, and the hypersonic algorithm for atmospheric parameter calculation, and the atmospheric parameters include the angle of attack, the sideslip angle, and the Mach number.

[0130] The input of the atmospheric parameter calculation is the pressure data of the pressure measurement points. The core calculation algorithms are divided into three parts: subsonic, transonic, and hypersonic. During use, the output parameters of other devices (such as inertial navigation devices) are used as references for mode switching to enable accurate calculation of atmospheric parameters at different flight speeds (speed ranges).

[0131] Referring to Table 1, the specific implementation method of the solution mode conversion is as follows:

[0132] Assume that the reference Mach number is represented as Ma d , Mad ≥ 0. Combining the Mach number ranges of different speed regimes and the Mach number range for the transition between two speed regimes, four Mach number ranges are determined as the basis for switching the solution modes. The first Mach number range is 0 ≤ Ma d < 0.6, the second Mach number range is 0.6 ≤ Ma d < 1, the third Mach number range is 1 ≤ Ma d < 4, and the fourth Mach number range is 4 ≤ Ma d < 6. It can be seen that the upper limit of the second Mach number range exceeds that of the first Mach number range, the upper limit of the third Mach number range exceeds that of the second Mach number range, and the upper limit of the fourth Mach number range exceeds that of the third Mach number range.

[0133] In specific implementation, determining the target solution mode of atmospheric parameters using the reference Mach number further includes:

[0134] 1) When 0 ≤ Ma d < 0.6, perform the subsonic algorithm to calculate the atmospheric parameters;

[0135] 2) When 0.6 ≤ Ma d < 1, simultaneously perform the subsonic algorithm and the transonic algorithm to calculate the atmospheric parameters, and take the solution result of the algorithm whose calculated flight speed (Mach number) is closer to the reference Mach number provided by the reference inertial navigation as the accurate calculated values of parameters such as angle of attack, sideslip angle, and Mach number;

[0136] 3) When 1 ≤ Ma d < 4, perform the transonic algorithm to calculate the atmospheric parameters;

[0137] 4) When 4 ≤ Ma d < 6, simultaneously perform the transonic algorithm and the hypersonic algorithm to calculate the atmospheric parameters, and take the solution result of the algorithm whose calculated flight speed (Mach number) is closer to the reference Mach number provided by the reference inertial navigation as the accurate calculated values of parameters such as angle of attack, sideslip angle, and Mach number;

[0138] 5) When Ma d ≥ 6, perform the hypersonic algorithm to calculate the atmospheric parameters.

[0139] Table 1 Algorithm Mode Status

[0140] Serial number Mad Subsonic algorithm Transonic algorithm Hypersonic algorithm 1 Ma0 to Ma0.6 On Off Off 2 Ma0.6 to Ma1 On On Off 3 Ma1 to Ma4 Off On Off 4 Ma4 to Ma0.6 Off On On 5 ≥Ma6 Off Off On

[0141] In some specific implementations, the principle of calculating atmospheric parameters provided in this embodiment is as Figure 3 shown.

[0142] The subsonic algorithm as a whole adopts a nested loop structure and uses two loop iterations. The small loop is the angle of attack and sideslip angle calculation loop, and the large loop is the Mach number calculation iteration. The subsonic algorithm is in the Figure 3 branch corresponding to module A. Further, the subsonic algorithm includes:

[0143] 1) Extract the pressure data of the third, fourth, fifth, eighth, and tenth pressure measurement points. Here, the pressure data of the corresponding pressure measurement points are represented by P3, P4, P5, P8, and P10.

[0144] 2) Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP3 of the third pressure measurement point, and the pressure coefficient CpsP8 of the eighth pressure measurement point. Cpa0 = (P3 - P8) / (P3 - P10), Cpb0 = (P4 - P5) / (P3 - P8), CpsP3 = (P3 - P s ) / (P t -P s ), CpsP8 = (P8 - P s ) / (P t -P s ). P3 represents the pressure data of the third pressure measurement point, P8 represents the pressure data of the eighth pressure measurement point, P10 represents the pressure data of the tenth pressure measurement point, P4 represents the pressure data of the fourth pressure measurement point, P5 represents the pressure data of the fifth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0145] 3) Set the initial Mach number within the first Mach number range and start the Mach number iterative solution loop;

[0146] 4) During the Mach number solution loop, iteratively solve the angle of attack and sideslip angle based on the pre-established subsonic basic coefficient table;

[0147] 5) When the calculated angle of attack and sideslip angle converge, solve the corresponding total pressure and static pressure according to the subsonic basic coefficient table and the current angle of attack and sideslip angle, and calculate the current Mach number using the solved total pressure and static pressure;

[0148] 6) When the calculated current Mach number converges, output the current angle of attack, sideslip angle, and Mach number.

[0149] The above-mentioned subsonic basic coefficient table is a coefficient table formed by normalizing the pressure data of each preset pressure measurement point at a given speed and attitude obtained from subsonic (Ma0 to Ma0.8) CFD simulation under subsonic conditions into dimensionless coefficients.

[0150] The overall transonic algorithm adopts a loop-nested structure and uses two loop iterations. The inner loop is the angle of attack and sideslip angle calculation loop, and the outer loop is the Mach number calculation iteration. The transonic algorithm is in the Figure 3 branch corresponding to module B. Further, the transonic algorithm includes:

[0151] 1) Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the ninth pressure measurement point;

[0152] 2) Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measurement point, and the pressure coefficient CpsP8 of the eighth pressure measurement point. Cpa0 = (P2 - P9) / (P1 - P8), Cpb0 = (P4 - P5) / (P1 - P8), CpsP1 = (P1 - P s ) / (P t -P s ), CpsP8 = (P8 - P s ) / (P t -P s ). P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P8 represents the pressure data of the eighth pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P4 represents the pressure data of the fourth pressure measurement point, P5 represents the pressure data of the fifth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0153] 3) Set the initial Mach number within the third Mach number range and start the Mach number iterative calculation loop;

[0154] 4) During the Mach number calculation loop, iteratively calculate the angle of attack and sideslip angle based on the pre-established transonic basic coefficient table;

[0155] 5) When the calculated angle of attack and sideslip angle converge, solve the corresponding total pressure and static pressure according to the transonic basic coefficient table and the current angle of attack and sideslip angle, and calculate the current Mach number using the solved total pressure and static pressure;

[0156] 6) When the calculated current Mach number converges, output the current angle of attack, sideslip angle, and Mach number.

[0157] The above-mentioned transonic basic coefficient table is a coefficient table formed by normalizing the pressure data of each preset pressure measurement point at a given speed and attitude obtained from transonic (Ma0.8 - Ma5) CFD simulation under transonic conditions into dimensionless coefficients.

[0158] Hypersonic flight is usually accompanied by altitudes significantly higher than those of conventional flights. Due to the thin air and low air pressure at high altitudes, altitude has an important impact on the calculation of hypersonic Mach numbers. Therefore, hypersonic algorithms take into account the influence of barometric altitude. The overall hypersonic algorithm adopts a nested loop structure and uses two loop iterations. The small loop is for iterative calculations of Mach number, angle of attack, and sideslip angle, and the large loop is for iterative calculations of barometric altitude. The hypersonic algorithm is branched in the Figure 3 corresponding module C of

[0159] 1) Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point;

[0160] 2) Construct the Mach number characteristic coefficient CpMa0, the angle of attack characteristic coefficient Cpa0, the sideslip angle characteristic coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measurement point, and the pressure coefficient CpsP9 of the ninth pressure measurement point. CpMa0 = P1 / P9, Cpa0 = (P1 - P9) / (P2 - P9), Cpb0 = (P6 - P7) / (P1 - P9), CpsP1 = (P1 - P s ) / (P t - P s ), CpsP9 = (P9 - P s ) / (P t - P s ). P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P6 represents the pressure data of the sixth pressure measurement point, P7 represents the pressure data of the seventh pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0161] 3) Set the initial barometric altitude and start the iterative solution loop for barometric altitude;

[0162] 4) During the barometric altitude solution loop, iteratively solve the angle of attack, sideslip angle, and Mach number based on a pre-established hypersonic basic coefficient table;

[0163] 5) When the solved angle of attack, sideslip angle, and Mach number converge, solve the corresponding total pressure and static pressure according to the hypersonic basic coefficient table and the current angle of attack, sideslip angle, and Mach number, and use the solved total pressure and static pressure to calculate the current barometric altitude;

[0164] 6) When the solved current barometric altitude converges, output the current angle of attack, sideslip angle, and Mach number.

[0165] The above hypersonic basic coefficient table is a coefficient table formed by obtaining the pressure data of each preset pressure measurement point at a given speed and attitude according to hypersonic (exceeding Ma5) CFD simulation under hypersonic conditions, and normalizing the pressure data obtained by the simulation into dimensionless coefficients.

[0166] Step 103, solve the atmospheric parameters in the target solution mode.

[0167] On the basis of determining the target solution mode, that is, determining whether to execute one algorithm or two algorithms. When performing the solution of atmospheric parameters using two algorithms among the subsonic algorithm, the transonic algorithm, and the hypersonic algorithm, after solving the atmospheric parameters in the target solution mode, it further includes:

[0168] Step 104, compare the Mach numbers in the two sets of atmospheric parameters obtained by the two algorithms with the reference Mach number; determine the set of atmospheric parameters corresponding to the Mach number with a relatively smaller difference from the reference Mach number as the solution result of the atmospheric parameters.

[0169] In specific implementation, when 0.6 ≤ Ma d <1 or 4 ≤ Ma d <6, two algorithms are simultaneously enabled to solve the atmospheric parameters to prevent the algorithm from failing due to too rapid speed change and inability to switch modes in a timely manner under high overload conditions of a waverider configuration wide - domain aircraft. In this case, two sets of output parameters will be output simultaneously: the subsonic algorithm solution and the transonic algorithm solution, or the transonic algorithm solution and the hypersonic algorithm solution.

[0170] For the case of multiple solutions, this embodiment designs a discrimination scheme. Judge the difference between the Mach number solution output values Ma A 、Ma B 、Ma C and the reference Mach number Ma d , and take the solution result with the Mach number solution output value closer to the reference Mach number Ma d as the final output parameter.

[0171] The method provided in this embodiment selects the pressure measurement points, performs the solution using one or two algorithms in different situations to obtain more accurate atmospheric parameters, and fully considers the problems that it is difficult to directly arrange sensors for a waverider configuration wide - domain aircraft during high - speed flight, and it is difficult to perform pressure measurement and parameter solution throughout the entire wide - speed range flight. The solution algorithm is constructed from three speed domains: subsonic, transonic, and hypersonic, forming an atmospheric parameter solution model for a wide - domain aircraft. This solution has high solution accuracy and fast solution speed.

[0172] Embodiment 2

[0173] Another embodiment of the present invention relates to a wide-speed-range atmospheric parameter calculation device. The implementation details of the wide-speed-range atmospheric parameter calculation device in this embodiment will be specifically described below. The following content is only the implementation details provided for convenient understanding and is not necessary for implementing this solution. The schematic diagram of the wide-speed-range atmospheric parameter calculation device in this embodiment can be as shown in Figure 4 shown, including:

[0174] A data acquisition module 201, configured to acquire the pressure data and reference Mach number of preset pressure measurement points on a wide-domain aircraft with a target waverider configuration.

[0175] A mode determination module 202, configured to determine the target calculation mode of atmospheric parameters by using the reference Mach number. The target calculation mode includes performing one or two of the subsonic algorithm, transonic algorithm, and hypersonic algorithm for calculating atmospheric parameters. The atmospheric parameters include angle of attack, sideslip angle, and Mach number.

[0176] A parameter calculation module 203, configured to calculate atmospheric parameters in the target calculation mode.

[0177] In a specific implementation, through CFD (Computational Fluid Dynamics) simulation, the pressure distribution on the surface of the wide-domain aircraft with the target waverider configuration under different flight states can be analyzed, and the points with better regular pressure changes on the surface are selected as the preset pressure measurement points, that is: the preset pressure measurement points are the pressure measurement points with regular pressure changes on the upper and lower surfaces of the wide-domain aircraft with the target waverider configuration.

[0178] In an example, the layout of the preset pressure measurement points is as shown in Figure 2 shown.

[0179] Figure 2 In the figure, the top view direction of the wide-domain aircraft with the target waverider configuration is shown. The preset pressure measurement points include the first pressure measurement point P1, the second pressure measurement point P2, the third pressure measurement point P3, the fourth pressure measurement point P4, the fifth pressure measurement point P5, the sixth pressure measurement point P6, the seventh pressure measurement point P7, and the eighth pressure measurement point P8 on the upper surface of the head of the wide-domain aircraft with the target waverider configuration, and the ninth pressure measurement point P9 and the tenth pressure measurement point P10 on the lower surface of the head of the fuselage of the wide-domain aircraft with the target waverider configuration.

[0180] Among them, the first pressure measurement point P1, the second pressure measurement point P2, the third pressure measurement point P3, the eighth pressure measurement point P8, and the ninth pressure measurement point P9 and the tenth pressure measurement point P10 are located on the intersection line of the vertical symmetry plane of the nose and the surface of the head of the fuselage. The fourth pressure measurement point P4 and the fifth pressure measurement point P5 are symmetric about the vertical symmetry plane of the nose. The sixth pressure measurement point P6 and the seventh pressure measurement point P7 are symmetric about the vertical symmetry plane of the nose.

[0181] Specifically, the pressure data of the third pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the tenth pressure measurement point are combined for the solution of atmospheric parameters in the subsonic algorithm:

[0182] 1) P3: Used as the upper surface pressure for the solution of the angle of attack pressure difference and as the large pressure point for the approximate dynamic pressure difference;

[0183] 2) P4 and P5: Mainly used for the solution of the sideslip angle;

[0184] 3) P8, used as the small pressure point for providing the approximate dynamic pressure difference;

[0185] 4) P10: Used as the lower surface pressure for the solution of the angle of attack pressure difference.

[0186] Specifically, the pressure data of the first pressure measurement point, the second pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the ninth pressure measurement point are combined for the solution of atmospheric parameters in the transonic algorithm:

[0187] 1) P1: Used as the large pressure point for the approximate dynamic pressure difference;

[0188] 2) P2: Used as the upper surface pressure for the solution of the angle of attack pressure difference;

[0189] 3) P4 and P5: Mainly used for the solution of the sideslip angle;

[0190] 4) P8: Used as the small pressure point for providing the approximate dynamic pressure difference;

[0191] 5) P9: Used as the lower surface pressure for the solution of the angle of attack pressure difference.

[0192] Specifically, the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point are combined for the solution of atmospheric parameters in the hypersonic algorithm:

[0193] 1) P1: Used as the large pressure point for the approximate dynamic pressure difference;

[0194] 2) P2: Used as the upper surface pressure for the solution of the angle of attack pressure difference;

[0195] 3) P4, P5, P6 and P7: Mainly used for the solution of the sideslip angle;

[0196] 4) P9: Used as the lower surface pressure for the solution of the angle of attack pressure difference and as the small pressure point for the approximate dynamic pressure difference.

[0197] In practical applications, the reference Mach number is determined by means of the output parameters of other devices, such as inertial navigation devices, to perform the solution mode switching so as to accurately solve the parameters at different flight speeds.

[0198] The input for atmospheric parameter calculation is the pressure data at the pressure measurement points. The core calculation algorithm is divided into three parts: subsonic, transonic, and hypersonic algorithms. During use, the output parameters of other devices (such as inertial navigation devices) are used as references for mode switching to enable accurate calculation of atmospheric parameters at different flight speeds (speed ranges).

[0199] Refer to Table 1. The specific implementation method of the calculation mode conversion is as follows:

[0200] Assume the reference Mach number is denoted as Ma d , M ad ≥0. Combining the Mach number ranges of different speed ranges and the Mach number ranges for transitions between two speed ranges, four Mach number ranges are determined as the basis for switching the calculation mode. The first Mach number range is 0 ≤ Ma d <0.6, the second Mach number range is 0.6 ≤ Ma d <1, the third Mach number range is 1 ≤ Ma d <4, and the fourth Mach number range is 4 ≤ Ma d <6. It can be seen that the upper limit of the second Mach number range exceeds the upper limit of the first Mach number range, the upper limit of the third Mach number range exceeds the upper limit of the second Mach number range, and the upper limit of the fourth Mach number range exceeds the upper limit of the third Mach number range.

[0201] In specific implementation, the target calculation mode of atmospheric parameters is determined using the reference Mach number, which further includes:

[0202] 1) When 0 ≤ Ma d <0.6, execute the subsonic algorithm to calculate atmospheric parameters;

[0203] 2) When 0.6 ≤ Ma d <1, execute both the subsonic algorithm and the transonic algorithm to calculate atmospheric parameters. Refer to the reference Mach number data provided by inertial navigation and take the calculation result of the algorithm whose calculated flight speed (Mach number) is closer to the reference Mach number as the accurate calculation value of parameters such as angle of attack, sideslip angle, and Mach number;

[0204] 3) When 1 ≤ Ma d <4, execute the transonic algorithm to calculate atmospheric parameters;

[0205] 4) When 4 ≤ Ma d <6, execute both the transonic algorithm and the hypersonic algorithm to calculate atmospheric parameters. Refer to the reference Mach number data provided by inertial navigation and take the calculation result of the algorithm whose calculated flight speed (Mach number) is closer to the reference Mach number as the accurate calculation value of parameters such as angle of attack, sideslip angle, and Mach number;

[0206] 5) When Ma dWhen it is ≥6, a hypersonic algorithm is executed to calculate the atmospheric parameters.

[0207] Table 1 Algorithm Mode Status

[0208]

[0209]

[0210] In some specific implementations, the principle of calculating the atmospheric parameters provided in this embodiment is as Figure 3 shown.

[0211] The subsonic algorithm as a whole adopts a nested loop structure and uses two loop iterations. The small loop is the loop for calculating the angle of attack and sideslip angle, and the large loop is the iteration for calculating the Mach number. The subsonic algorithm corresponds to the branch of module A in Figure 3 . Further, the subsonic algorithm includes:

[0212] 1) Extract the pressure data of the third pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point, and the tenth pressure measurement point. Here, the pressure data of the corresponding pressure measurement points are represented by P3, P4, P5, P8, and P10.

[0213] 2) Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP3 of the third pressure measurement point, and the pressure coefficient CpsP8 of the eighth pressure measurement point. Cpa0 = (P3 - P8) / (P3 - P10), Cpb0 = (P4 - P5) / (P3 - P8), CpsP3 = (P3 - P s ) / (P t - P s ), CpsP8 = (P8 - P s ) / (P t - P s ). P3 represents the pressure data of the third pressure measurement point, P8 represents the pressure data of the eighth pressure measurement point, P10 represents the pressure data of the tenth pressure measurement point, P4 represents the pressure data of the fourth pressure measurement point, P5 represents the pressure data of the fifth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0214] 3) Set an initial Mach number within the first Mach number range and start the Mach number iterative calculation loop;

[0215] 4) During the Mach number calculation loop, iteratively calculate the angle of attack and sideslip angle based on a pre-established subsonic basic coefficient table;

[0216] 5) When the calculated angle of attack and sideslip angle converge, according to the subsonic basic coefficient table and the current angle of attack and sideslip angle, solve the corresponding total pressure and static pressure, and use the solved total pressure and static pressure to calculate the current Mach number;

[0217] 6) When the calculated current Mach number converges, output the current angle of attack, sideslip angle and Mach number.

[0218] The above-mentioned subsonic basic coefficient table is a coefficient table formed by normalizing the pressure data of each preset pressure measurement point at a given speed and attitude obtained from subsonic (Ma0 to Ma0.8) CFD simulation into dimensionless coefficients.

[0219] The whole transonic algorithm adopts a loop-nested structure and uses two loop iterations. The small loop is the angle of attack and sideslip angle calculation loop, and the large loop is the Mach number calculation iteration. The transonic algorithm is in Figure 3 the branch corresponding to module B, and further, the transonic algorithm includes:

[0220] 1) Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the fourth pressure measurement point, the fifth pressure measurement point, the eighth pressure measurement point and the ninth pressure measurement point;

[0221] 2) Construct the initial angle of attack pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measurement point and the pressure coefficient CpsP8 of the eighth pressure measurement point, Cpa0 = (P2 - P9) / (P1 - P8), Cpb0 = (P4 - P5) / (P1 - P8), CpsP1 = (P1 - P s ) / (P t - P s )、CpsP8 = (P8 - P s ) / (P t - P s ), P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P8 represents the pressure data of the eighth pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P4 represents the pressure data of the fourth pressure measurement point, P5 represents the pressure data of the fifth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0222] 3) Set the initial Mach number within the third Mach number range and start the Mach number iterative calculation loop;

[0223] 4) During the Mach number calculation loop, iteratively calculate the angle of attack and sideslip angle based on the pre-established transonic basic coefficient table;

[0224] 5) When the calculated angle of attack and sideslip angle converge, according to the transonic basic coefficient table and the current angle of attack and sideslip angle, solve the corresponding total pressure and static pressure, and calculate the current Mach number using the solved total pressure and static pressure;

[0225] 6) When the calculated current Mach number converges, output the current angle of attack, sideslip angle, and Mach number.

[0226] The above-mentioned transonic basic coefficient table is a coefficient table formed by normalizing the pressure data of each preset pressure measurement point at a given speed and attitude obtained from transonic (Ma0.8 - Ma5) CFD simulation into dimensionless coefficients according to the transonic condition.

[0227] Hypersonic flight is usually accompanied by a significantly higher altitude than conventional flight. Due to the thin air and low air pressure at high altitudes, altitude has an important impact on the calculation of hypersonic Mach numbers. Therefore, the hypersonic algorithm takes into account the influence of air pressure altitude. The overall hypersonic algorithm adopts a nested loop structure and uses two loop iterations. The small loop is for iterative calculations of Mach number, angle of attack, and sideslip angle, and the large loop is for iterative calculations of air pressure altitude. The hypersonic algorithm is in Figure 3 the branch corresponding to module C. Further, the hypersonic algorithm includes:

[0228] 1) Extract the pressure data of the first pressure measurement point, the second pressure measurement point, the sixth pressure measurement point, the seventh pressure measurement point, and the ninth pressure measurement point;

[0229] 2) Construct the Mach number characteristic coefficient CpMa0, the angle of attack characteristic coefficient Cpa0, the sideslip angle characteristic coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measurement point, and the pressure coefficient CpsP9 of the ninth pressure measurement point. CpMa0 = P1 / P9, Cpa0 = (P1 - P9) / (P2 - P9), Cpb0 = (P6 - P7) / (P1 - P9), CpsP1 = (P1 - P s ) / (P t -P s ),CpsP9 = (P9 - P s ) / (P t -P s ),where P1 represents the pressure data of the first pressure measurement point, P2 represents the pressure data of the second pressure measurement point, P6 represents the pressure data of the sixth pressure measurement point, P7 represents the pressure data of the seventh pressure measurement point, P9 represents the pressure data of the ninth pressure measurement point, P t represents the total pressure, and P s represents the static pressure;

[0230] 3) Set the initial air pressure altitude and start the iterative calculation loop of air pressure altitude;

[0231] 4) During the air pressure altitude calculation loop, iteratively calculate the angle of attack, sideslip angle, and Mach number based on a pre-established hypersonic basic coefficient table;

[0232] 5) When the calculated angle of attack, sideslip angle, and Mach number converge, solve the corresponding total pressure and static pressure according to the hypersonic basic coefficient table and the current angle of attack, sideslip angle, and Mach number, and use the solved total pressure and static pressure to calculate the current air pressure altitude;

[0233] 6) When the calculated current air pressure altitude converges, output the current angle of attack, sideslip angle, and Mach number.

[0234] The above-mentioned hypersonic basic coefficient table is a coefficient table formed by normalizing the pressure data of each preset pressure measurement point at a given speed and attitude obtained from hypersonic (exceeding Ma5) CFD simulation under hypersonic conditions into dimensionless coefficients.

[0235] On the basis of determining the target calculation mode, that is, determining whether to execute one algorithm or two algorithms. When performing atmospheric parameter calculation using two algorithms among the subsonic algorithm, transonic algorithm, and hypersonic algorithm, after calculating the atmospheric parameters in the target calculation mode, it further includes:

[0236] In some implementation manners, the device of this embodiment further includes:

[0237] A result determination module 204, configured to compare the Mach numbers in the two sets of atmospheric parameters calculated by the two sets of algorithms with a reference Mach number; and determine the set of atmospheric parameters corresponding to the Mach number with a relatively smaller difference from the reference Mach number as the result of the atmospheric parameter calculation.

[0238] In a specific implementation, when 0.6 ≤ Ma d <1 or 4 ≤ Ma d <6, two algorithms are simultaneously enabled to calculate the atmospheric parameters to prevent the algorithm from failing due to too rapid speed change and inability to switch modes in a timely manner under high overload conditions of a waverider configuration wide-domain aircraft. In this case, two sets of output parameters will be output simultaneously: the subsonic algorithm solution and the transonic algorithm solution, or the transonic algorithm solution and the hypersonic algorithm solution.

[0239] For the case of multiple solutions, this embodiment designs a discrimination scheme. Judge the magnitude of the difference between the Mach number calculation output values Ma A 、Ma B 、Ma C and the reference Mach number Ma d , and take the solution result with the Mach number calculation output value closer to the reference Mach number Ma d as the final output parameter.

[0240] In this embodiment, the pressure measurement points are selected, and one or two algorithms are executed under different conditions for calculation to obtain more accurate atmospheric parameters. Considering that it is difficult to directly arrange sensors for the waverider configuration wide-range aircraft during high-speed flight, and due to the difficulty of performing pressure measurement and parameter calculation throughout the wide-speed range flight, a calculation algorithm is constructed from three speed ranges: subsonic, transonic, and hypersonic, forming a wide-range aircraft atmospheric parameter calculation model. This solution has high calculation accuracy and a fast calculation speed.

[0241] It is worth mentioning that each module involved in this embodiment is a logical module. In practical applications, a logical unit can be a physical unit, a part of a physical unit, or can be implemented as a combination of multiple physical units. In addition, to highlight the innovative part of the present invention, units that are not closely related to solving the technical problems proposed by the present invention are not introduced in this embodiment, but this does not mean that there are no other units in this embodiment.

[0242] Embodiment 3

[0243] Another embodiment of the present invention relates to an electronic device, including:

[0244] At least one processor; and,

[0245] A memory communicatively connected to the at least one processor; wherein,

[0246] The memory stores instructions executable by the at least one processor, and the instructions are executed by the at least one processor to enable the at least one processor to execute the wide-speed range atmospheric parameter calculation method of the above embodiment.

[0247] Among them, the memory and the processor are connected by a bus. The bus can include any number of interconnected buses and bridges, and the bus connects various circuits of one or more processors and the memory together. The bus can also connect various other circuits such as peripheral devices, voltage regulators, and power management circuits, which are well known in the art, so they will not be further described herein. The bus interface provides an interface between the bus and the transceiver. The transceiver can be an element or multiple elements, such as multiple receivers and transmitters, providing units for communicating with various other devices on the transmission medium. The data processed by the processor is transmitted over the wireless medium through the antenna. Further, the antenna also receives data and transmits the data to the processor.

[0248] The processor is responsible for managing the bus and normal processing, and can also provide various functions, including timing, peripheral interface, voltage regulation, power management, and other control functions. The memory can be used to store the data used by the processor when performing operations.

[0249] Embodiment 4

[0250] Another embodiment of the present invention relates to a computer-readable storage medium storing a computer program. When the computer program is executed by a processor, the above method embodiments are implemented.

[0251] Those skilled in the art can understand that all or part of the steps in implementing the above method embodiments can be completed by instructing relevant hardware through a program. The program is stored in a storage medium, including several instructions for causing a device (which can be a single-chip microcomputer, a chip, etc.) or a processor to execute all or part of the steps of the methods described in the various embodiments of the present invention. The foregoing storage medium includes: various media such as USB flash drives, mobile hard disks, read-only memories (ROM), random access memories (RAM), magnetic disks, or optical discs that can store program codes.

[0252] This embodiment also provides a computer program product, including a computer program, which when executed by a processor, implements the steps of the wide-speed-range atmospheric parameter calculation method in the above embodiment.

[0253] Embodiment 5

[0254] This embodiment provides an application example, which provides specific implementation manners of subsonic algorithms, transonic algorithms, and hypersonic algorithms.

[0255] The specific implementation manner of the subsonic algorithm model is as follows:

[0256] 1) Extract the pressure data at points such as P3, P4, P5, P8, and P10.

[0257] 2) Use the data at three position points such as P3, P8, and P10 to construct the initial angle of attack differential pressure coefficient Cpa0.

[0258] Cpa0 = (P3 - P8) / (P3 - P10);

[0259] 3) Use the data at four position points such as P3, P4, P5, and P8 to construct the initial sideslip angle differential pressure coefficient Cpb0.

[0260] Cpb0 = (P4 - P5) / (P3 - P8);

[0261] 4) Establish the pressure coefficient CpsP3 of P3 and the pressure coefficient CpsP8 of P8.

[0262] CpsP3 = (P3 - P s ) / (Pt -P s )、CpsP8 = (P8 - P s ) / (P t -P s ),P t is the total pressure, and P s is the static pressure;

[0263] 5) Set the initial Mach number Ma0 = 0.4, and interpolate in the subsonic basic coefficient table to obtain the angle of attack pressure difference coefficient table CPa_ma0, the sideslip angle pressure difference coefficient table CPb_ma0, and the pressure coefficient tables CpsP3_ma0 and CpsP8_ma0 of P3 and P8 corresponding to the initial Mach number.

[0264] Among them, the subsonic basic coefficient table is formed by obtaining the surface pressure values at each pressure measurement point under the given speed and attitude through subsonic (Ma0 to Ma0.8) CFD simulation and normalizing them into dimensionless coefficients. The basic coefficient table contains four variables: pressure coefficient, Mach number, angle of attack, and sideslip angle;

[0265] 6) Set the initial sideslip angle β (0,0) = 0, substitute Cpa0 and β (0,0) into CPa_ma0, and interpolate to obtain the corresponding angle of attack α (0,1) ;

[0266] 7) Substitute Cpb0 and α (0,1) into CPb_ma0, and interpolate to obtain the corresponding sideslip angle β (0,1) ;

[0267] 8) Replace β (0,1) with β (0,0) and repeat steps 6) and 7), and iteratively calculate the angle of attack and the sideslip angle until convergence.

[0268] From the perspective of the algorithm loop, a more general description is adopted. Replace β (n,m+1) with β (n,m+1) and repeat steps 6) and 7), and iteratively calculate until both the angle of attack and the sideslip angle converge, that is, the convergence condition is satisfied:

[0269] |β(n,m + 1) - β(n,m)| < △1 && |α(n,m + 1) - α(n,m)| < △2,

[0270] where △1 and △2 are both convergence factors, and α (n,m) , β (n,m) represent the output values of the angle of attack and sideslip angle loops when the Mach number loop is performed n times and the attack sideslip angle loop is performed m times. n and m represent the number of times of the Mach number loop and the angle of attack and sideslip angle loop respectively, and n and m are both greater than or equal to 0;

[0271] If the convergence condition is satisfied, output α(n,m + 1) and β(n,m + 1).

[0272] 9) Based on the obtained angle of attack and sideslip angle, interpolate in the CpsP3_ma0 and CpsP8_ma0 tables to obtain CpsP3’_ma0 and CpsP8'_ma0, establish the system of binary linear equations CpsP3’_ma0 = (P3 - Ps1) / (Pt1 - Ps1), CpsP8'_ma0 = (P8 - Ps1) / (Pt1 - Ps1), and solve the equations simultaneously for Pt1 and Ps1; and calculate Ma1 according to the aforementioned equations (13) and (14).

[0273] 10) Replace Ma0 with Ma1, and repeat steps 5), 6), 7), 8), and 9) until convergence. From the perspective of the algorithm loop, a more general description is adopted. Replace Ma n+1 with Ma n Repeat steps 5), 6), 7), 8), and 9), and iteratively calculate the Mach number until convergence, that is, |Ma n+1 - Ma n | < △3, where △3 is the convergence factor. Here, Ma n represents the output value of the Mach number after n cycles. Finally, output parameters such as the angle of attack αA, sideslip angle β A , Mach number Ma A , etc.

[0274] The specific implementation method of the transonic algorithm is as follows:

[0275] 1) Extract the pressure data of pressure measurement points such as P1, P2, P4, P5, P8, and P10.

[0276] 2) Use the pressure data of 4 pressure measurement points, namely P1, P2, P8, and P9, to construct the initial angle of attack differential pressure coefficient Cpa0, where Cpa0 = (P2 - P9) / (P1 - P8).

[0277] 3) Use the pressure data of 4 pressure measurement points, namely P1, P4, P5, P8, to construct the initial sideslip angle differential pressure coefficient Cpb0, where Cpb0 = (P4 - P5) / (P1 - P8).

[0278] 4) Establish the pressure coefficient CpsP1 of P1 and the pressure coefficient CpsP8 of P8, where CpsP1 = (P1 - P s ) / (P t - P s ), CpsP8 = (P8 - P s ) / (P t - P s ), Pt is the total pressure, and Ps is the static pressure.

[0279] 5) Set the initial Mach number Ma0 = 2, and interpolate in the transonic basic coefficient table to obtain the angle of attack pressure difference coefficient table CPa_ma0, the sideslip angle pressure difference coefficient table CPb_ma0, and the pressure coefficient tables CpsP1_ma0 and CpsP8_ma0 of P1 and P8 corresponding to the initial Mach number. Among them, the transonic basic coefficient table is obtained by CFD simulation of transonic (Ma0.8 - Ma5) to get the surface pressure values at each pressure measurement point under given speed and attitude, and normalize them into dimensionless coefficients to form a coefficient table. The basic coefficient table contains four variables: pressure coefficient, Mach number, angle of attack, and sideslip angle;

[0280] 6) Set the initial sideslip angle β (0,0) = 0, substitute Cpa0 and β (0,0) into CPa_ma0, and interpolate to obtain the corresponding angle of attack α (0,1) ;

[0281] 7) Substitute Cpb0 and α (0,1) into CPb_ma0, and interpolate to obtain the corresponding sideslip angle β (0,1) ;

[0282] 8) Replace β (0,1) with β (0,0) and repeat steps 6) and 7), iteratively calculate the angle of attack and sideslip angle until convergence. From the perspective of algorithm loop, a more general description is adopted. Replace β (n,m+1) with β (n,m+1) and repeat steps 6) and 7), iteratively calculate until both the angle of attack and sideslip angle converge, that is, |β(n,m + 1)-β(n,m)| < △1 && |α(n,m + 1)-α(n,m)| < △2, where △1 and △2 are both convergence factors, then output α(n,m + 1) and β(n,m + 1). Among them, α (n,m) and β (n,m) represent the loop output values of the angle of attack and sideslip angle when the Mach number loop is performed n times and the attack sideslip angle loop is performed m times. n and m respectively represent the number of times of the Mach number loop and the angle of attack and sideslip angle loop, and n, m are both greater than or equal to 0;

[0283] 9) According to the obtained angle of attack and sideslip angle, interpolate in the CpsP1_ma0 and CpsP8_ma0 tables to obtain CpsP1’_ma0 and CpsP8'_ma0, establish a system of binary linear equations CpsP1’_ma0 = (P1 - Ps1) / (Pt1 - Ps1), CpsP8'_ma0 = (P8 - Ps1) / (Pt1 - Ps1), solve Pt1 and Ps1 simultaneously; and calculate Ma1 according to the aforementioned formulas (13) and (14);

[0284] 10) Replace Ma0 with Ma1, and repeat steps 5), 6), 7), 8), and 9) until convergence. From the perspective of the algorithm loop, a more general description is adopted. Replace Ma n+1 with Ma n Repeat steps 5), 6), 7), 8), and 9), and iteratively calculate the Mach number until convergence, that is, |Ma n+1 - Ma n | < △3, where △3 is the convergence factor. Here, Ma n represents the output value of the Mach number after n loops. Finally, output the angle of attack αB, sideslip angle β B , Mach number Ma B and other parameters.

[0285] The specific implementation method of the hypersonic algorithm is as follows:

[0286] 1) Extract the pressure data of the pressure measurement points such as P1, P2, P6, P7, and P9;

[0287] 2) Use the pressure data of 2 pressure measurement points such as P1 and P9 to construct the Ma characteristic coefficient CpMa0, where CpMa0 = P1 / P9;

[0288] 3) Use the pressure data of 3 pressure measurement points such as P1, P2, and P9 to construct the angle of attack characteristic coefficient Cpa0, where Cpa0 = (P1 - P9) / (P2 - P9);

[0289] 4) Use the pressure data of 6 pressure measurement points such as P1, P6, P7, and P9 to construct the sideslip angle characteristic coefficient Cpb0, where Cpb0 = (P6 - P7) / (P1 - P9);

[0290] 5) Construct the pressure coefficients of P1 and P9, CpsP1 = (P1 - P s ) / (P t - P s ), CpsP9 = (P9 - P s ) / (P t - P s ), where P t is the total pressure and P s is the static pressure.

[0291] 6) Set the initial barometric altitude \(H_{p0} = 30000m\), and interpolate in the hypersonic basic coefficient table to obtain the \(Ma\) coefficient table \(C_{pMa\_Hp0}\), the angle of attack pressure difference coefficient table \(C_{pa\_Hp0}\), the sideslip angle pressure difference coefficient table \(C_{pb\_Hp0}\), and the pressure coefficient tables \(C_{psP1\_Hp0}\) and \(C_{psP9\_Hp0}\) of \(P1\) and \(P9\) corresponding to \(H_{p0}\). Among them, according to hypersonic (exceeding \(Ma5\)) CFD simulations, the surface pressure values at each pressure measurement point are obtained at a given speed and attitude, and are normalized into dimensionless coefficients to form a coefficient table. The hypersonic basic coefficient table contains five variables: barometric altitude, pressure coefficient, Mach number, angle of attack, and sideslip angle;

[0292] 7) Set the initial sideslip angle \(\beta\) (0,0) \(= 0\), substitute \(\beta\) (0,0) and the initial values of \(C_{pa0}\) and \(C_{pMa0}\) into \(C_{pa\_Hp0}\) and \(C_{pMa\_Hp0}\) respectively, and interpolate to obtain the corresponding angle of attack \(\alpha\) (0,1) and \(Ma\) (0,1) ;

[0293] 8) Substitute \(C_{pb0}\), \(\alpha\) (0,1) and \(Ma\) (0,1) into \(C_{Pb\_Hp0}\), and interpolate to obtain the corresponding sideslip angle \(\beta\) (0,1) ;

[0294] 9) Replace \(\beta\) (0,1) with \(\beta\) (0,0) and repeat steps 7 and 8, and iteratively calculate the angle of attack, sideslip angle, and Mach number until convergence. From the perspective of the algorithm loop, a more general description is adopted. Replace \(\beta\) (n,m+1) with \(\beta\) (n,m+1) and repeat steps 7 and 8, and iteratively calculate until the angle of attack, sideslip angle, and Mach number all converge, that is, satisfy the convergence:

[0295] \(|\beta(n,m + 1)-\beta(n,m)|<\triangle1\&\&|\alpha(n,m + 1)-\alpha(n,m)|<\triangle2\&\&|Ma(n,m + 1)-Ma(n,m)|<\triangle3\),

[0296] where \(\triangle1\), \(\triangle2\), and \(\triangle3\) are all convergence factors, and \(\alpha\) (n,m) and \(\beta\) (n,m) and \(Ma(n,m)\) respectively represent the Mach number, angle of attack, and sideslip angle cyclic output values when the barometric altitude cycle is carried out \(n\) times, and the Mach number, angle of attack, and sideslip angle cycles are carried out \(m\) times. \(n\) and \(m\) respectively represent the barometric altitude cycle and the Mach number, angle of attack, and sideslip angle cycle times, and \(n\) and \(m\) are both greater than or equal to 0;

[0297] If the convergence condition is satisfied, output α(n,m+1), β(n,m+1), and Ma(n,m+1). Among them, according to the obtained angle of attack, sideslip angle, and Mach number, interpolate in CpsP1 and CpsP9 to obtain CpsP1'_Hp0 and CpsP9'_Hp0 respectively, and establish a system of binary linear equations CpsP1'_Hp0 = (P1 - P s ) / (P t -P s ), CpsP9'_Hp0 = (P9 - P s ) / (P t -P s ), and solve the equations simultaneously for P t and P s ;

[0298] 10) According to the solved P t and P s , calculate the barometric altitude Hp1 according to Equation (15);

[0299] 11) Replace Hp0 with Hp1 and repeat steps 6), 7), 8), 9), 10), and 11). From the perspective of algorithm loop, a more general description is adopted. Replace Hp n+1 with Hp n and repeat steps 6), 7), 8), 9), 10), and 11), and iteratively calculate the barometric altitude until convergence, that is, |Hp n+1 - Hp n | < △4, where △4 is the convergence factor. Among them, Hp n represents the output value of the barometric altitude after n cycles. Finally, output parameters such as the angle of attack αC, sideslip angle β C , and Mach number Ma C .

[0300] When 0.6 ≤ Ma d < 1 or 4 ≤ Ma d < 6, two algorithms are simultaneously enabled for atmospheric parameter calculation to prevent the algorithm from failing due to the inability to switch modes in a timely manner due to too rapid speed changes under high overload conditions of the aircraft. In this case, the algorithm will output two sets of output parameters: subsonic algorithm solution and transonic / supersonic algorithm solution, or transonic / supersonic algorithm solution and hypersonic algorithm solution.

[0301] For the case of multiple solutions, the algorithm model designs a discrimination scheme. Judge the difference between the Mach number calculation output values Ma A , Ma B , Ma C and the reference Mach number Ma d , and take the calculation result of the Mach number calculation output value that is closer to the reference Mach number Ma d as the final output parameter.

[0302] On the basis of constructing the foregoing algorithm, an embedded atmospheric data system is implemented based on the wide-speed-range atmospheric parameter calculation method. Further, multiple calibration tests are carried out to obtain higher parameter calculation accuracy during actual use. The following calibration methods can be referred to:

[0303] 1) The process of calibrating the embedded atmospheric data system based on ground pressure excitation, that is, the hardware-in-the-loop simulation test;

[0304] 2) The process of obtaining excitation condition calibration by hanging and flying other aircraft with the help of a scaled model, that is, the flight test;

[0305] 3) The process of obtaining excitation condition calibration by using a wind tunnel with the help of a scaled model, that is, the wind tunnel test.

[0306] Through methods such as hardware-in-the-loop simulation tests, flight tests, and wind tunnel tests, the calculation accuracy of this method can be verified from different verification types (static / dynamic), different excitation methods (pressure / airflow excitation), and different speed ranges. Through the test data, the original CFD simulation-based data table can be corrected and updated, so that the algorithm model has higher calculation accuracy.

[0307] Those of ordinary skill in the art can understand that the above embodiments are specific embodiments for implementing the present invention, and in actual applications, various changes can be made to it in form and details without departing from the spirit and scope of the present invention.

Claims

1. A method for calculating atmospheric parameters in a wide speed range, characterized in that: include: Obtain pressure data and reference Mach number of preset pressure measuring points on the target waverider configuration wide-range aircraft; Determine a target solution mode of atmospheric parameters by using the reference Mach number, wherein the target solution mode comprises executing one or two algorithms of a subsonic algorithm, a transonic algorithm and a hypersonic algorithm to solve atmospheric parameters, wherein the atmospheric parameters comprise an angle of attack, a sideslip angle and a Mach number; The atmospheric parameters are solved in the target solving mode.

2. The wide-speed range atmospheric parameter calculation method according to claim 1, characterized in that: The preset pressure measuring points are pressure measuring points where the pressure on the upper and lower surfaces of the target waverider configuration wide-range aircraft changes regularly.

3. The wide-speed range atmospheric parameter calculation method according to claim 1, characterized in that: The target solution mode of determining the atmospheric parameters by using the reference Mach number includes: When the reference Mach number is within a first Mach number range, a subsonic algorithm is executed; when the reference Mach number is within a second Mach number range, a subsonic algorithm and a transonic algorithm are executed simultaneously; when the reference Mach number is within a third Mach number range, a transonic algorithm is executed; when the reference Mach number is within a fourth Mach number range, a transonic algorithm and a hypersonic algorithm are executed simultaneously; when the reference Mach number is greater than an upper limit of the fourth Mach number range, a hypersonic algorithm is executed; Among them, the second Mach number range exceeds the upper limit of the first Mach number range, the third Mach number range exceeds the upper limit of the second Mach number range, and the fourth Mach number range exceeds the upper limit of the third Mach number range.

4. The wide-speed range atmospheric parameter calculation method according to claim 3, characterized in that: The preset pressure measuring points include a first pressure measuring point, a second pressure measuring point, a third pressure measuring point, a fourth pressure measuring point, a fifth pressure measuring point, a sixth pressure measuring point, a seventh pressure measuring point and an eighth pressure measuring point located on the upper surface of the head of the target waverider configuration wide-range aircraft, and a ninth pressure measuring point and a tenth pressure measuring point located on the lower surface of the fuselage head of the target waverider configuration wide-range aircraft; The pressure data combination of the third pressure measuring point, the fourth pressure measuring point, the fifth pressure measuring point, the eighth pressure measuring point and the tenth pressure measuring point is used for solving the atmospheric parameters of the subsonic algorithm, the pressure data combination of the first pressure measuring point, the second pressure measuring point, the fourth pressure measuring point, the fifth pressure measuring point, the eighth pressure measuring point and the ninth pressure measuring point is used for solving the atmospheric parameters of the transonic algorithm, and the pressure data combination of the first pressure measuring point, the second pressure measuring point, the sixth pressure measuring point, the seventh pressure measuring point and the ninth pressure measuring point is used for solving the atmospheric parameters of the hypersonic algorithm.

5. The wide-speed range atmospheric parameter calculation method according to claim 4, characterized in that: The subsonic algorithm includes: extracting pressure data of the third pressure measuring point, the fourth pressure measuring point, the fifth pressure measuring point, the eighth pressure measuring point and the tenth pressure measuring point; Construct the initial attack angle pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient of the third pressure measuring point CpsP3 and the pressure coefficient of the eighth pressure measuring point CpsP8, Cpa0 = (P3-P8) / (P3-P10), Cpb0 = (P4-P5) / (P3-P8), CpsP3 = (P3-P s ) / (P t -P s ), CpsP8=(P8-P s ) / (P t -P s ), P3 represents the pressure data of the third pressure measuring point, P8 represents the pressure data of the eighth pressure measuring point, P10 represents the pressure data of the tenth pressure measuring point, P4 represents the pressure data of the fourth pressure measuring point, P5 represents the pressure data of the fifth pressure measuring point, P t Indicates total pressure, P s Indicates static pressure; An initial Mach number is set within the first Mach number range, and a Mach number iterative solution cycle is started; In the Mach number solution cycle, the angle of attack and sideslip angle are iteratively solved based on the pre-established subsonic basic coefficient table; When the calculated angle of attack and sideslip angle converge, the corresponding total pressure and static pressure are solved according to the subsonic basic coefficient table and the current angle of attack and sideslip angle, and the current Mach number is solved using the solved total pressure and static pressure; When the calculated current Mach number converges, the current angle of attack, sideslip angle and Mach number are output.

6. The wide-speed range atmospheric parameter calculation method according to claim 4, characterized in that: The transonic algorithm includes: Extracting pressure data of the first pressure measuring point, the second pressure measuring point, the fourth pressure measuring point, the fifth pressure measuring point, the eighth pressure measuring point and the ninth pressure measuring point; Construct the initial attack angle pressure difference coefficient Cpa0, the initial sideslip angle pressure difference coefficient Cpb0, the pressure coefficient of the first pressure measuring point CpsP1 and the pressure coefficient of the eighth pressure measuring point CpsP8, Cpa0 = (P2-P9) / (P1-P8), Cpb0 = (P4-P5) / (P1-P8), CpsP1 = (P1-P s ) / (P t -P s ), CpsP8=(P8-P s ) / (P t -P s ), P1 represents the pressure data of the first pressure measuring point, P2 represents the pressure data of the second pressure measuring point, P8 represents the pressure data of the eighth pressure measuring point, P9 represents the pressure data of the ninth pressure measuring point, P4 represents the pressure data of the fourth pressure measuring point, P5 represents the pressure data of the fifth pressure measuring point, P t Indicates total pressure, P s Indicates static pressure; An initial Mach number is set within the third Mach number range, and a Mach number iterative solution cycle is started; During the Mach number solution cycle, the angle of attack and sideslip angle are iteratively solved based on the pre-established transonic basic coefficient table; When the calculated angle of attack and sideslip angle converge, the corresponding total pressure and static pressure are solved according to the transonic basic coefficient table and the current angle of attack and sideslip angle, and the current Mach number is solved using the solved total pressure and static pressure; When the calculated current Mach number converges, the current angle of attack, sideslip angle and Mach number are output.

7. The wide-speed range atmospheric parameter calculation method according to claim 4, characterized in that: The hypersonic algorithm includes: Extracting pressure data of the first pressure measuring point, the second pressure measuring point, the sixth pressure measuring point, the seventh pressure measuring point and the ninth pressure measuring point; Construct the Mach number characteristic coefficient CpMa0, the angle of attack characteristic coefficient Cpa0, the sideslip angle characteristic coefficient Cpb0, the pressure coefficient CpsP1 of the first pressure measuring point and the pressure coefficient CpsP9 of the ninth pressure measuring point, CpMa0 = P1 / P9, Cpa0 = (P1-P9) / (P2-P9), Cpb0 = (P6-P7) / (P1-P9), CpsP1 = (P1-P s ) / (P t -P s ), CpsP9=(P9-P s ) / (P t -P s ), P1 represents the pressure data of the first pressure measuring point, P2 represents the pressure data of the second pressure measuring point, P6 represents the pressure data of the sixth pressure measuring point, P7 represents the pressure data of the seventh pressure measuring point, P9 represents the pressure data of the ninth pressure measuring point, P t Indicates total pressure, P s Indicates static pressure; Set the initial pressure altitude and start the pressure altitude iterative solution cycle; During the pressure altitude solution cycle, the angle of attack, sideslip angle and Mach number are iteratively solved based on the pre-established hypersonic basic coefficient table; When the angle of attack, sideslip angle and Mach number obtained by the solution converge, the corresponding total pressure and static pressure are solved according to the hypersonic basic coefficient table and the current angle of attack, sideslip angle and Mach number, and the current pressure altitude is solved using the solved total pressure and static pressure; When the calculated current pressure altitude converges, the current angle of attack, sideslip angle and Mach number are output.

8. The wide-speed range atmospheric parameter calculation method according to any one of claims 5 to 7, characterized in that: The subsonic basic coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measuring point at a given speed and attitude under subsonic speed, and normalizing the pressure data obtained by simulation into dimensionless coefficients; The transonic basic coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measuring point at a given speed and attitude under transonic speed, and normalizing the pressure data obtained by simulation into dimensionless coefficients; The hypersonic basic coefficient table is a coefficient table formed by simulating the pressure data of each preset pressure measuring point at a given speed and attitude under hypersonic speed, and normalizing the pressure data obtained by simulation into dimensionless coefficients.

9. The wide-speed range atmospheric parameter calculation method according to claim 1, characterized in that: In the case of executing two algorithms among the subsonic algorithm, the transonic algorithm and the hypersonic algorithm to solve the atmospheric parameters, after solving the atmospheric parameters in the target solving mode, the method further includes: The Mach numbers in the two sets of atmospheric parameters solved by the two sets of algorithms are compared with the reference Mach numbers; A set of atmospheric parameters corresponding to a Mach number with a relatively small difference from a reference Mach number is determined as the atmospheric parameter solution result.

10. An electronic device, characterized in that: include: at least one processor; as well as, a memory communicatively connected to the at least one processor; wherein, The memory stores instructions that can be executed by the at least one processor, and the instructions are executed by the at least one processor so that the at least one processor can execute the wide-speed domain atmospheric parameter solution method as described in any one of claims 1 to 9.

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