Method for optimizing and determining cruise flight range performance of hypersonic flight vehicle

Abstract

The invention belongs to the technical field of hypersonic flight vehicle cruise flight range performance determination, and particularly relates to a hypersonic flight vehicle cruise flight range performance optimization determination method, which comprises the following steps: step 1, calculating an effective mass correction coefficient of an aircraft, where V is the speed of the aircraft, g is the gravitational acceleration, and r is the radius of the earth; 2, calculating the maximum lift-drag ratio of the aircraft; 3, calculating the fuel weight coefficient epsilon of the aircraft; and 4, calculating the optimal voyage of the aircraft, wherein Isp is the specific impulse of the aircraft.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of cruise flight range performance determination of hypersonic aircraft, and particularly relates to a cruise flight range performance optimization determination method of hypersonic aircraft. BACKGROUND

[0002] The hypersonic aircraft refers to an aircraft with a flight speed exceeding 5 Mach. The range calculation of the hypersonic aircraft is affected by various factors, such as flight altitude, speed, aerodynamic characteristics, fuel consumption rate, etc.

[0003] The existing range calculation methods, such as the calculation method based on the Breguet formula, are mainly applicable to low-speed aircraft. Under the condition of hypersonic speed, the applicability is limited due to the influence of factors such as centrifugal effect and change of aerodynamic characteristics.

[0004] During the cruise process of the hypersonic aircraft, energy conservation is the core law of its flight. However, the existing range calculation methods for high-speed aircraft usually ignore the influence of the change of the speed and the lift-drag ratio of the aircraft on the range, and do not consider the dynamic relationship between the normalized lift coefficient and the fuel weight coefficient, resulting in that the range calculation result is not accurate enough and it is difficult to meet the optimization design requirements of the hypersonic aircraft.

[0005] Therefore, there is an urgent need for a hypersonic aircraft range optimization calculation method which can comprehensively consider the factors such as the aerodynamic characteristics of the aircraft, fuel consumption and speed change, so as to provide effective support for the optimization design of the hypersonic aircraft. SUMMARY

[0006] The purpose of the application is to provide a cruise flight range performance optimization determination method of hypersonic aircraft, which combines the law of energy conservation and the principle of aerodynamics, considers the normalized lift coefficient and the cruise fuel weight coefficient, and gives a range calculation model, so as to improve the accuracy of the range calculation of the hypersonic aircraft and provide effective support for the optimization design of the hypersonic aircraft.

[0007] The technical scheme of the application is:

[0008] A cruise flight range performance optimization determination method of hypersonic aircraft, comprising:

[0009] Step one, calculating the effective mass correction coefficient of the aircraft Wherein, V is the speed of the aircraft, g is the gravitational acceleration, and r is the radius of the earth.

[0010] Step two, calculating the maximum lift-drag ratio of the aircraft

[0011] Step three, calculating the fuel weight coefficient ε of the aircraft

[0012] Step four, calculating the optimal range of the aircraft

[0013]

[0014] wherein,

[0015] Isp is the specific impulse of the aircraft.

[0016] According to at least one embodiment of the present application, in the above-mentioned cruise flight range performance optimization determination method of the hypersonic aircraft, the gravity acceleration g is 9.81 m / s^2;

[0017] The earth radius r is 6371000 m.

[0018] According to at least one embodiment of the present application, in the above-mentioned cruise flight range performance optimization determination method of the hypersonic aircraft, the maximum lift-drag ratio of the aircraft is calculated Specifically,

[0019]

[0020] wherein, k is a dimensionless constant related to the shape of the aircraft, C D0 is the zero-lift drag coefficient of the aircraft.

[0021] According to at least one embodiment of the present application, in the above-mentioned cruise flight range performance optimization determination method of the hypersonic aircraft, the dimensionless constant k related to the shape of the aircraft is 1.

[0022] According to at least one embodiment of the present application, in the above-mentioned cruise flight range performance optimization determination method of the hypersonic aircraft, the fuel weight coefficient ε of the aircraft is calculated, specifically:

[0023]

[0024] wherein, W initial is the initial weight of the aircraft, and W final is the final weight of the aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0025] Figure 1 is a schematic diagram of the cruise flight trajectory of the aircraft provided by the embodiments of the present application;

[0026] Figure 2 is a schematic diagram of the cruise flight range performance optimization determination method of the hypersonic aircraft provided by the embodiments of the present application;

[0027] Figure 3 is a relationship curve between the cruise flight range R and the cruise flight fuel weight coefficient in the cruise flight phase provided by the embodiments of the present application;

[0028] Figure 4 is a relationship curve between the range R and the specific impulse Isp of the aircraft in a cruise flight phase provided by the embodiment of the present application.

[0029] In order to better illustrate the embodiments, some contents in the drawings are omitted, enlarged or reduced, which are only used for exemplary illustration and cannot be understood as a limitation to the present application. DETAILED DESCRIPTION

[0030] In order to make the technical solutions of the present application and the advantages thereof clearer, the technical solutions of the present application will be further clearly and completely described below with reference to the drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application, which are used to explain the present application, but not to limit the present application. It should be noted that, in order to facilitate the description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design.

[0031] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of the present application should be the general meanings understood by the general skilled in the art. In the description of the present application, "comprising" indicates that the concept appearing before the word covers the concepts listed after the word and its equivalents, and does not exclude other associated concepts.

[0032] The cruise flight trajectory of the aircraft is shown in Figure 1 As shown in the figure, in the cruise process of the aircraft, the energy conservation is a basic physical law, and the total energy is composed of potential energy and kinetic energy, and the change rate of the total energy is equal to the net thrust power. According to the law of conservation of energy, the change rate of the total energy can be expressed as:

[0033]

[0034] The left side of formula (1) is is the kinetic energy of the aircraft, which depends on the mass and flight speed of the aircraft, and the left side of mgh is the gravitational potential energy of the aircraft, which is related to the mass, gravitational acceleration and flight height of the aircraft.

[0035] The right side of formula (1) (T-D)·V represents the net thrust power, that is, the work done by the resultant force of the thrust and the resistance in unit time.

[0036] Comparing the change rate of the total energy with the net thrust power, when the lift is equal to the gravity, the total energy expression can be arranged as:

[0037]

[0038] For a low-speed aircraft, assuming that the lift is equal to the resistance, and the speed and specific energy are considered as constants, the Breguet range formula is obtained:

[0039]

[0040] For hypersonic vehicles, whose flight speed is high enough to make centrifugal effect important, the effective weight of the hypersonic vehicle will be corrected as:

[0041]

[0042]

[0043]

[0044]

[0045] That is:

[0046]

[0047] There are two cases to be discussed:

[0048] Case 1: When the speed and the lift-drag ratio remain unchanged, the range is:

[0049]

[0050] Case 2: When the speed remains unchanged and the lift-drag ratio changes, the lift-drag ratio expression is:

[0051]

[0052] The drag can be decomposed into zero-lift drag C D0 and lift-induced drag The lift-drag ratio expression is:

[0053]

[0054] The derivative of C L with respect to the angle of attack is taken for equation (10), and:

[0055]

[0056] We get:

[0057]

[0058]

[0059] The lift coefficient when the maximum lift-drag ratio is obtained is:

[0060]

[0061]

[0062] According to the hypersonic tangent ogive theory, the lift coefficient and the maximum lift-drag ratio at the maximum lift-drag ratio are respectively:

[0063]

[0064]

[0065] It is obtained:

[0066]

[0067]

[0068] That is:

[0069]

[0070] Bring formula (18) into formula (7), there is:

[0071]

[0072]

[0073]

[0074]

[0075]

[0076] That is:

[0077]

[0078] Because:

[0079]

[0080] There is:

[0081]

[0082] It can be obtained:

[0083]

[0084] Bring formula (23) and formula (25) into formula (22), there is:

[0085]

[0086]

[0087]

[0088]

[0089] That is,

[0090]

[0091] Further simplifying equation (24) gives:

[0092]

[0093] Let We have:

[0094]

[0095]

[0096]

[0097]

[0098] Substitute equation (26) and equation (27) into equation (25):

[0099]

[0100] Because:

[0101]

[0102] Substitute equation (31) into equation (30):

[0103]

[0104]

[0105] That is,

[0106]

[0107] Substitute equation (28) and equation (29) into equation (33):

[0108]

[0109] Substitute equation (16) into equation (34):

[0110]

[0111] Equation (35) reflects the relationship between the cruising range of the aircraft and the weight, the related parameters of the lift-drag ratio, the dynamic pressure, etc., where h is the height, H is the scale height, H = 8500 m, and p0 is the sea level atmospheric density. However, the form of equation (35) is relatively complex. In order to more clearly study the effects of aerodynamics and fuel on the range, a normalized lift coefficient Ω is introduced.

[0112]

[0113] When the hypersonic vehicle is a slender vehicle, n = 2, equation (16) and equation (24) are substituted into equation (36), and then:

[0114]

[0115]

[0116] Let:

[0117]

[0118]

[0119] Substituting equation (37) into equation (35), we have:

[0120]

[0121] Because the fuel weight coefficient ε during the cruise phase can be expressed as:

[0122]

[0123]

[0124] Transforming, we have:

[0125] z = y (1 - ε) = Ω (1 - ε) … … (40)

[0126] According to the trigonometric transformation:

[0127]

[0128] Substituting equation (41) into equation (38), we have:

[0129]

[0130] Using Taylor series expansion, we have:

[0131]

[0132] Substituting into equation (43), and taking the first three terms, we have:

[0133]

[0134]

[0135] Substitute formula (44) into formula (42), we have:

[0136]

[0137] Derivate formula (42) with respect to Ω and set it equal to zero, we have:

[0138]

[0139] Because Formula (46) can be changed to:

[0140]

[0141] According to the derivative of composite function, formula (47) can be changed to:

[0142]

[0143] The optimal Ω is obtained from formula (48):

[0144] Ω opt = (1-ε) -1 / 2 ……(49)

[0145] Substitute formula (49) into formula (42) and formula (45) respectively, the optimal range R can be obtained:

[0146]

[0147]

[0148] Formula (50) and formula (51) can provide a theoretical basis for the optimization of cruise flight range.

[0149] Based on the above, given the initial parameters of the hypersonic vehicle in the cruise phase, including flight speed V, specific impulse Isp, maximum lift-drag ratio Lift coefficient C L , zero-lift-drag coefficient C D0 , reference area S ref , sea level atmospheric density ρ0, scale height H and gravitational acceleration g, vehicle shape related dimensionless constant k, initial weight W initial , final weight W final , dynamic pressure q ∞ , Earth radius r, the calculation of the optimal range of the hypersonic vehicle can be referred to as follows.

[0150] Calculate the fuel weight coefficient of the vehicle The ratio of fuel consumption to initial weight.

[0151] Optimize the normalized lift coefficient, find the optimal normalized lift coefficient Ω that maximizes the range by taking the derivative and setting it to zero opt = (1-ε) -1 / 2 .

[0152] Calculate the maximum lift-drag ratio For slender hypersonic vehicles, k can be taken as 1.

[0153] Substitute the optimized Ω into the range formula to calculate the optimal range opt

[0154] Based on the above, the application provides a hypersonic vehicle cruise flight range performance optimization determination method, as shown in Figure 1 Based on the law of conservation of energy, and considering the influence of the change of lift coefficient and fuel weight on the range.

[0155] Step 1, calculate the effective mass correction coefficient of the aircraft Where V is the speed of the aircraft, g is the acceleration of gravity, and r is the radius of the earth.

[0156] The acceleration of gravity g can be taken as 9.81 m / s^2, and the radius of the earth r can be taken as 6371000 m.

[0157] Step 2, calculate the maximum lift-drag ratio of the aircraft

[0158]

[0159] Where k is a dimensionless constant related to the shape of the aircraft, C D0 is the zero-lift-drag coefficient of the aircraft.

[0160] The dimensionless constant k related to the shape of the aircraft can be taken as 1.

[0161] Step 3, calculate the fuel weight coefficient ε of the aircraft:

[0162]

[0163] Where W initial is the initial weight of the aircraft, and W final is the final weight of the aircraft.

[0164] Step 4, calculate the optimal range of the aircraft

[0165]

[0166] Where,​

[0167] Isp is the specific impulse of the aircraft.

[0168] In a specific example, the speed of the aircraft V = 3000 m / s, the specific impulse Isp = 1200 s, the maximum lift-drag ratio Initial weight W initial = 10000 kg, the final weight W final = 8000 kg, the cruise flight range performance optimization determination method of the hypersonic aircraft disclosed in the above embodiment is implemented as follows.

[0169] Calculate the fuel weight coefficient of the aircraft

[0170] Optimize the normalized lift coefficient, find the optimal normalized lift coefficient Ω when maximizing the range by taking the derivative and setting it to zero opt = (1-0.2) -1 / 2 = 1.118.

[0171] Calculate the maximum lift-drag ratio of the aircraft

[0172] Substitute the optimized Ω opt into the range formula to calculate the optimal range:

[0173]

[0174] Figure 3 is the relationship curve between the cruise flight range R of the aircraft and the fuel weight coefficient during the cruise flight phase. The cruise speed is 3000 m / s, the specific impulse Isp is 1200 s, and the maximum lift-drag ratio is 3.54.

[0175] Figure 4 is the relationship curve between the cruise flight range R of the aircraft and the specific impulse Isp during the cruise flight phase. The cruise speed is 3000 m / s, the fuel weight coefficient is 0.2, and the maximum lift-drag ratio is 3.54.

[0176] The cruise flight range performance optimization determination method of the hypersonic aircraft disclosed in the above embodiment comprehensively considers various key factors involved in the cruise flight process of the hypersonic aircraft, including flight speed, flight altitude, thrust, drag, fuel consumption rate, and aerodynamic characteristics.

[0177] The existing method simplifies the complex factors when calculating the range of a hypersonic vehicle, which results in an incomplete and inaccurate calculation result. The hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment can more comprehensively reflect the range performance of the vehicle by combining the law of conservation of energy and the principle of aerodynamics, and comprehensively analyzing the power characteristics, aerodynamic characteristics and fuel consumption law of the vehicle. The hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment breaks through the limitations of the existing method and is suitable for range calculation under more complex flight conditions.

[0178] The hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment introduces normalized lift coefficient and cruise fuel weight coefficient, which significantly improves the accuracy of range calculation. The normalized lift coefficient takes into account the dynamic characteristics of the vehicle under hypersonic conditions, and can more accurately reflect the relationship between lift and drag. The cruise fuel weight coefficient takes into account the influence of initial and final fuel weight, avoiding the simplification in fuel consumption calculation of traditional methods. By introducing these parameters, the accuracy of range calculation under complex flight conditions can be significantly improved, especially for range optimization calculation of hypersonic vehicles. This method can provide more reliable theoretical support for the design of hypersonic vehicles, and is suitable for vehicle selection, design and performance evaluation.

[0179] During the design phase of the vehicle, engineers can use the hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment to calculate the range of the vehicle, so as to evaluate the range performance of different design schemes and select the optimal power system or aerodynamic shape. During the flight mission planning phase, the hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment can be used to calculate the range of the vehicle, which can help to evaluate the feasibility of different flight paths and optimize the flight scheme to improve mission efficiency.

[0180] The hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiment has clear steps and intuitive results in actual calculation. Through explicit mathematical expressions and optimization steps, the complexity and uncertainty in the calculation process can be reduced, the calculation efficiency can be improved, and the method is easy for engineers to apply. The method is suitable for different types of hypersonic vehicles and has wide applicability. Whether it is for slender body vehicles or other types of hypersonic vehicles, the range optimization calculation can be realized by adjusting the relevant parameters, and the method has application flexibility and universality.

[0181] The hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiments introduces new variables and parameters, such as normalized lift coefficient and cruise fuel weight coefficient, to construct a more refined and perfect range calculation model. This method not only has innovation in theory, but also shows advanced technical level in practical application. Compared with existing methods, it can more accurately reflect the dynamics characteristics of the hypersonic vehicle and its influence on the range, and provides a new idea and method for the design and optimization of the vehicle.

[0182] In summary, the hypersonic vehicle cruise flight range performance optimization determination method disclosed in the above embodiments has significant superiority in comprehensiveness, accuracy, practicality, calculation efficiency and engineering application, and can provide strong theoretical support and practical tools for the design, optimization and performance evaluation of the hypersonic vehicle, and has broad application prospect and far-reaching academic influence.

[0183] So far, the technical solution of the present application has been described in combination with the preferred embodiments shown in the drawings, and those skilled in the art should understand that the protection scope of the present application is obviously not limited to these specific embodiments, and those skilled in the art can make equivalent changes or replacements to the related technical features without deviating from the principles of the present application, and the technical solutions after these changes or replacements will fall within the protection scope of the present application.

Claims

1. A method for optimizing and determining the cruise flight range performance of a hypersonic vehicle, characterized in that, Comprises: Step one, calculate the effective mass correction coefficient of the aircraft Wherein, V is the speed of the aircraft, g is the acceleration of gravity, r is the earth radius; Step two, calculating the maximum lift-to-drag ratio of the aircraft Step three, calculate the fuel weight coefficient ε of the aircraft; Step four, calculating the optimal range of the aircraft Wherein, Isp is the specific impulse of the aircraft.

2. The method of claim 1, wherein, The gravity acceleration g takes 9.81 m / s^2; The earth radius r takes 6371000 m.

3. The method of claim 2, wherein, Calculating maximum lift-drag ratio of an aircraft Specifically: where k is a non-dimensional constant related to the shape of the aircraft, C D0 is the zero-lift drag coefficient of the aircraft.

4. The method of claim 3, wherein, The aircraft shape related dimensionless constant k takes 1.

5. The method of claim 4, wherein, Calculate the fuel weight coefficient ε of the aircraft, specifically: where W initial is the initial weight of the aircraft, W final is the final weight of the aircraft.

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