Overall parameter closed-loop design method for vertical launching aircraft

Abstract

The invention discloses a closed-loop design method for overall parameters of a vertical launch aircraft, which comprises the following steps: determining design indexes of the vertical launch aircraft, and giving design parameters of the vertical launch aircraft; setting initial values of appearance parameters of the vertical launching aircraft for iterative calculation, and starting iterative calculation: determining aerodynamic characteristic parameters, determining fuel relative mass of the vertical launching aircraft by solving a relative quantity motion differential equation set, and determining a fuel mass coefficient of a booster; calculating a fuel oil relative mass coefficient of a turbojet engine of the vertical launching aircraft; solving a wing load and a thrust-weight ratio; determining an empty aircraft weight coefficient; calculating the total take-off weight and the total initial weight of a patrol flight section of the vertical launch aircraft; calculating the fuselage length of the vertical launch aircraft; determining wing positions and rudder wing sizes; and repeating the iterative calculation process on the appearance parameters of the vertical launching aircraft until the iterative residual error of each appearance parameter is smaller than a preset requirement, and outputting a final appearance parameter.

Description

Technical Field

[0001] The present invention belongs to the technical field of overall design of aircraft, and in particular relates to a closed-loop design method for overall parameters of a vertical launch aircraft. Background Art

[0002] With technological advancements in the aircraft field, various projects have increasingly required ground-based vertical launch systems. Typical scenarios include solid rocket-assisted launches, turbojet-powered cruise missiles, and vertically launched missiles co-located on a platform. Currently, there is no comprehensive parameter closed-loop design method for the overall design of such vertical launch vehicles, and research is limited. Typically, a multidisciplinary optimization design approach is conducted based on a specific base vehicle, gradually meeting or improving performance requirements through iterative optimization. This design method is suitable for applications with a pre-existing base vehicle and a sufficient product portfolio. However, its iteration efficiency is low during preliminary design demonstrations, making it difficult to meet the requirements for multiple-alternative comparisons and parameter sensitivity analysis. A comprehensive parameter closed-loop design method uses theoretical calculations or semi-empirical models to obtain design results that directly meet the input requirements. Each design parameter and result are closed-loop and self-consistent within the theoretical model, eliminating the need for optimization. The advantages of closed-loop design methods include the availability of theoretical models, rapid design speed, and the ease of multiple-alternative comparisons and parameter sensitivity analysis during preliminary demonstrations. It also provides a relatively stable starting point for optimization. Therefore, a comprehensive parameter closed-loop design method is needed for vertical launch vehicles.

[0003] In terms of closed-loop parameter design, conventional missiles and aircraft currently have mature overall parameter closed-loop design methods. These methods typically use theoretical formulas to calculate the required missile or aircraft parameters from the required indicators. However, missiles and aircraft differ in their propulsion systems: missiles belong to the category of spacecraft, while aircraft belong to the category of aircraft. Consequently, their overall parameter closed-loop design methods differ significantly. For example, when evaluating fuel quality, the typical parameter of solid (or liquid) rocket engines commonly used in spacecraft is specific impulse, while the typical parameter of air-breathing propulsion systems used in aircraft is fuel consumption rate. The theoretical models for these two propulsion systems are quite different. Regarding profile design, spacecraft require a trajectory model, while aircraft require a flight performance model, resulting in significant differences in profile design approaches. For vertical launch vehicles (VLVs) using solid rocket engines for boost and turbojet engines for endurance, the initial phase exhibits typical VLV flight characteristics, while the cruise phase exhibits characteristics of long-duration aircraft flight. During initial design, aircraft design methods are typically used to first design the cruise phase. Based on the cruise phase design results, the booster design is then evaluated. This involves two iterations in parameter design, resulting in a long cycle and low efficiency. Summary of the Invention

[0004] The purpose of the present invention is to provide a closed-loop design method for the overall parameters of a vertical launch vehicle. Based on given indicator inputs and design parameters, calculations are carried out through established theoretical models, semi-empirical models, etc., to obtain design results that can meet the indicator requirements, and the theoretical relationship between each parameter can be closed-loop and self-consistent.

[0005] In order to achieve the above tasks, the present invention adopts the following technical solutions:

[0006] A closed-loop design method for overall parameters of a vertical launch vehicle comprises:

[0007] Determine the design indicators of the vertical launch vehicle and give the design parameters of the vertical launch vehicle;

[0008] Set the initial values ​​of the vertical launch vehicle's shape parameters for iterative calculation and start the iterative calculation:

[0009] Determine aerodynamic characteristic parameters based on the parametric shape of the vertical launch vehicle;

[0010] Combined with the design parameters, the relative mass of the fuel of the vertical launch vehicle is determined by solving the differential equations of relative motion, and the fuel mass coefficient of the booster is determined;

[0011] Calculating the relative fuel mass coefficient of the turbojet engine of the vertical launch vehicle based on the design indicators and design parameters;

[0012] Based on the design indicators, solve the wing loading and thrust-to-weight ratio;

[0013] Calculate the empty aircraft weight coefficient through the relative mass of fuel and the relative mass coefficient of oil;

[0014] Calculate the takeoff gross weight and cruising phase starting gross weight of the vertical launch vehicle using the empty mass coefficient, fuel mass coefficient, and fuel mass coefficient in combination with design indicators.

[0015] Using the design parameters, calculate the fuselage length of the vertical launch vehicle;

[0016] Determine the wing position and rudder wing size based on the design indicators and design parameters;

[0017] Repeat the above iterative calculation process for the shape parameters of the vertical launch vehicle until the iterative residual of each shape parameter is less than the preset requirement, and output the final shape parameters.

[0018] Furthermore, the external parameters of the vertical launch vehicle include fuselage length, wingspan, chord length, wing position, rudder wingspan length, takeoff gross weight, cruising phase starting gross weight, and wing loading.

[0019] Furthermore, based on the parameterized shape of the vertical launch vehicle, aerodynamic characteristic parameters are determined, including:

[0020] Based on the given wing installation angle / sweep angle / taper ratio, airfoil, and initial value of the shape parameters, the shape parameters are iteratively updated. The relevant dimensional parameters of the aircraft are described through formatting commands, and the grid is divided for calculation. After the calculation is completed, the fuselage drag model is used to correct the aerodynamic characteristic parameters including the lift coefficient and drag coefficient.

[0021] Furthermore, the relative mass of the fuel of the vertical launch vehicle is determined by solving the relative mass motion differential equations in combination with the design parameters, and the booster fuel mass coefficient is determined, including:

[0022] The relative quantity motion differential equations are constructed by the relative quantity method and solved to obtain the curves of velocity and trajectory angle varying with relative mass coefficient.

[0023] The interpolation method is used to intercept the relative mass coefficient when the trajectory inclination angle is less than the trajectory inclination angle at the end of the boost in the design parameters and the speed is greater than the speed at the end of the boost in the design parameters, and this relative mass coefficient is used as the booster fuel mass coefficient.

[0024] Furthermore, based on the design indicators and design parameters, the relative mass coefficient of fuel of the turbojet engine of the vertical launch vehicle is calculated, including:

[0025] The flight profile of a vertical launch vehicle can be divided into the launch adjustment phase, the cruise phase, and the strike phase. Based on the mass of the vertical launch vehicle at launch, the mass after the boost ends, the mass when entering stable cruise, the mass during the terminal mission, and the mass at mission completion, an expression for the relative mass coefficient of fuel is constructed.

[0026] The unknown quantity in the expression of the relative fuel mass coefficient is the ratio of the mass at mission completion to the mass at launch, which can be expressed as the product of the mass ratios of each mission segment. For the calculation of the mass ratio of the cruise segment, one state, either cruise or hover, is selected for design, or designs are performed for both states separately, and the design result with the larger relative fuel mass coefficient is taken. The mass ratio of the cruise segment is characterized with reference to the Breguet range-time formula. The fuel consumption rate used in the calculation is converted to the fuel consumption rate at the speed, Mach number, and altitude under cruise conditions based on the maximum thrust of the turbojet engine at sea level and in a stationary state and its corresponding fuel consumption rate.

[0027] Furthermore, based on the design indicators, the wing loading and thrust-to-weight ratio are solved, including:

[0028] Calculate the wing load for each operating condition based on the minimum speed, static ceiling, maximum normal circling overload, and maximum instantaneous overload required in the design indicators, and take the maximum value as the wing load;

[0029] According to the maximum flight speed and maximum climb angle required in the design indicators, the thrust-to-weight ratio for each operating condition is calculated, and the maximum value is taken as the thrust-to-weight ratio for the cruise phase and converted to the standard condition of zero altitude.

[0030] Furthermore, the empty weight coefficient is calculated using the relative fuel mass and the relative fuel mass coefficient, including:

[0031] The empty weight coefficient is calculated by combining the structural weight coefficient, the booster engine case weight coefficient, the fuel tank weight coefficient, and the turbojet engine weight coefficient, with the ratio of the structural weight to the cruise total weight, the ratio of the booster fuel weight to the booster total weight, the ratio of the fuel to the fuel tank total weight, the turbojet engine single-machine thrust-to-weight ratio, and the relative mass of the fuel and the relative mass coefficient of the fuel.

[0032] Furthermore, the fuselage length of the vertical launch vehicle is calculated using the design parameters, including:

[0033] Based on the fuel tank length coefficient, booster length coefficient, and turbojet engine fuselage length coefficient, the fuselage cross-sectional area is calculated in combination with the fuselage width / height of the step design parameters, and the fuselage length is calculated by adding the instrument compartment length in the design parameters.

[0034] Furthermore, the wing position and rudder wing size are determined based on the design indicators and design parameters, including:

[0035] First, the chord length of the wing is calculated using the aircraft's takeoff gross weight, wing loading, and the wing aspect ratio in the design index, and the wingspan is determined.

[0036] Secondly, the lift line slope of the airfoil is calculated using the aerodynamic efficiency, aspect ratio, and Mach number. The slope of the moment curve is determined by combining the lift line slope of the airfoil with the static stability in the design index. Combined with the control wall in the design index, the control efficiency of the rudder is calculated.

[0037] Finally, the rudder lift is calculated using the rudder control efficiency, center of mass position, wing chord length, and fuselage length; the rudder area is determined using the rudder lift and the slope of the lift line of the rudder wing; the rudder span and chord length are calculated from the rudder area and the rudder aspect ratio.

[0038] Furthermore, after the initial values ​​of the shape parameters are given, a set of values ​​of the shape parameters after each iterative calculation can be obtained, and the next iterative calculation is performed based on this. After obtaining the value of each shape parameter after the next iteration, the difference between the values ​​after this iteration is made to obtain the residual of the two adjacent iterations of each shape parameter; the iteration is repeated until the iterative residuals of all shape parameters are less than the specified value, and the parameter closed-loop design is completed.

[0039] A terminal device includes a processor, a memory, and a computer program stored in the memory; when the processor executes the computer program, the vertical launch aircraft overall parameter closed-loop design method is implemented.

[0040] A computer-readable storage medium stores a computer program; when the computer program is executed by a processor, the vertical launch vehicle overall parameter closed-loop design method is implemented.

[0041] Compared with the prior art, the present invention has the following technical features:

[0042] The method of the present invention can be applied to the overall design of vertical launch aircraft with solid rocket engine boost and turbojet engine endurance. It combines the differences and characteristics of aerospace and aviation aircraft designs and proposes an aircraft overall design calculation model that takes into account both vertical launch and cruise flight. The method creates a booster engine design calculation model based on relative quantity differential motion equations by giving index inputs and design parameters. The vertical turning section and the reconnaissance and cruise flight section are effectively combined through wing load, thrust-to-weight ratio and turning trajectory parameters. After modeling and calculation, a design result that can meet the index requirements can be obtained, ensuring that the theoretical relationship between each parameter can be closed-loop and self-consistent. The efficiency of parameter closed-loop design is effectively improved and the design cycle is reduced. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] Figure 1 Schematic diagram of the process of the present invention;

[0044] Figure 2 Schematic diagram showing the change of flight speed and trajectory inclination with the relative mass coefficient of fuel in an embodiment of the present invention;

[0045] Figure 3 : is a curve showing the static stability change with the random wing position in the embodiment of the present invention. DETAILED DESCRIPTION

[0046] In order to solve the problems existing in the existing parameter closed-loop design method, it is necessary to integrate the design methods of spacecraft and aircraft in the overall parameter closed-loop design, and conduct unified parameter design for the vertical launch vehicle as a whole (including the vertical launch stage and the cruise stage). In view of the characteristics of the vertical launch vehicle using solid rocket engines to assist vertical turns in the vertical launch stage and turbojet engines for endurance in the cruise stage, a vertical launch vehicle overall parameter closed-loop design method is provided, as follows:

[0047] (1) Determine the design parameters of the vertical launch vehicle.

[0048] Straight-and-level flight performance: maximum flight speed, minimum flight speed, static ceiling;

[0049] Fixed straight climbing performance: maximum climbing angle;

[0050] Endurance performance: flight time (long-term performance), range (long-distance performance);

[0051] Maneuverability: maximum normal circling overload, maximum instantaneous overload;

[0052] Carrying capacity: payload mass;

[0053] Maneuverability and stability: control ratio, static stability;

[0054] (2) Given the design parameters of the vertical launch vehicle.

[0055] Aerodynamic layout parameters: wing aspect ratio R l2c (or span), wing installation angle / sweep angle / taper ratio, rudder span (or aspect ratio R l2c-wd ), fuselage width / height (or fuselage diameter), wing / rudder airfoil;

[0056] Structural parameters: center of mass position (proportion of fuselage length) x cg , structural quality coefficient R me1 , Instrument cabin (payload cabin) length L load ;

[0057] Booster engine parameters: specific impulse I s , thrust-to-weight ratio R T2Wv or thrust;

[0058] Turbojet engine parameters: fuel consumption rate SFC, turbojet engine thrust-to-weight ratio R T2W ;

[0059] Fuel system parameters: residual oil coefficient K fuel ;

[0060] Ballistic parameters: vertical turn angle α, speed v at the end of boost, and ballistic inclination θ (or boost engine thrust, thrust time).

[0061] (3) Set the initial values ​​of the shape parameters of the vertical launch vehicle used for iterative calculation.

[0062] The initial values ​​of the vertical launch vehicle's fuselage length, wingspan, chord length, wing position (ratio of fuselage length), rudder wingspan, takeoff gross weight, cruising phase initial gross weight, and wing loading are determined. The initial values ​​only determine the starting point of the iterative calculations in steps (4) to (11) and do not affect the iterative convergence results.

[0063] (4) Determine the aerodynamic characteristic parameters based on the parametric shape of the vertical launch vehicle.

[0064] Establish a parametric shape of the vertical launch aircraft, and use existing mature aerodynamic calculation software such as the vortex grid method, the panel method or the computational fluid dynamics method to calculate the aerodynamic characteristic parameters. Taking the AVL calculation as an example, according to the given wing installation angle / sweep angle / taper ratio, airfoil, and the initial value of the shape parameters proposed in step (3) or the shape parameters updated iteratively in step (12), the aircraft-related size parameters are described by formatting commands, and the grid is divided for calculation. After the calculation is completed, the fuselage drag model such as the neural network fitting model is used to correct the aerodynamic characteristic parameters of the fuselage, including the lift coefficient C L , drag coefficient C D .

[0065] (5) Combined with the design parameters, the relative mass of the fuel of the vertical launch vehicle is determined by solving the relative mass motion differential equations, and the booster fuel mass coefficient is determined.

[0066] To calculate the parameters of solid rocket motors using the relative quantity method, it is necessary to solve the following relative quantity motion differential equations:

[0067]

[0068] The relative mass coefficient μ represents the relative mass of fuel consumed by the aircraft at time t (i.e., the ratio of consumed fuel to takeoff total weight), I s is the specific impulse; α is the angle of attack used during vertical turning, which can be given as a fixed value; v is the speed; R T2Wv is the thrust-to-weight ratio in the vertical turning section (i.e. the ratio of the boost thrust to the total takeoff weight); R W2S is the wing loading; θ is the ballistic inclination angle; x is the horizontal position, y is the vertical position; ρ is the air density, and g is the acceleration due to gravity.

[0069] Under given working conditions and initial values, the relative motion differential equations can be solved to obtain the curves of velocity v and trajectory inclination θ with μ. Generally, as μ increases, velocity v increases and trajectory inclination θ decreases. According to the velocity and trajectory inclination at the end of the boost given in step (2), the interpolation method is used to intercept the μ value when the trajectory inclination θ is less than the trajectory inclination at the end of the boost in step (2) and the velocity v is greater than the velocity at the end of the boost in step (2) (e.g. Figure 2 As shown), the μ value is used as the booster fuel mass coefficient R mfv .

[0070] (6) Based on the design indicators and design parameters, calculate the relative mass coefficient of fuel of the turbojet engine of the vertical launch vehicle.

[0071] The flight profile of an aircraft can be divided into the launch adjustment phase, the cruise phase, and the strike phase. Let the mass of the aircraft at launch be m0, the mass after the boost is m1, the mass when entering stable cruise is m2, the mass when performing the terminal mission is m3, and the mass when the mission is completed is m4. Let the mass of fuel consumed at the end of the mission be K of the total fuel mass. fuel times (i.e., there is residual oil or dead oil), then the relative mass coefficient of fuel R mfc It can be expressed as:

[0072]

[0073] Where m1 / m0 is the available booster fuel mass coefficient R mfv Indicates that m1 / m0=1-R mfv ; and m4 / m0 can be expressed as the product of the mass ratios of each task segment:

[0074]

[0075] Then we only need to obtain the mass ratio m of the i-th task segment and the i-1-th task segment i / m i-1 , the fuel relative mass coefficient R can be obtained mfc . According to engineering experience, the mass ratio of the launch adjustment section is m2 / m1=0.97, and the mass ratio of the strike section is m4 / m3=0.992; the mass ratio m3 / m2 of the cruise section needs to be calculated based on the range and flight time indicators in step (1). Choose one state of cruise (long-distance state) or hovering (long-term flight state) for design, or design for the two states separately and take the design result with the larger fuel mass; refer to Breguet's range-time formula, the mass ratio of this mission section determined by the range or flight time is:

[0076]

[0077] Among them, R r For the voyage, E t is the flight time, SFC is the fuel consumption rate, C L / C D is the lift-to-drag ratio in long-distance or long-duration flight state, and V is the cruising speed. For jet-powered aircraft, the lift-to-drag ratio in long-duration flight state is usually taken as the maximum lift-to-drag ratio of the aircraft, while the lift-to-drag ratio in long-distance flight state is 0.866 times the maximum lift-to-drag ratio.

[0078] The SFC used in solving the mass ratio m3 / m2 in the cruise phase needs to be based on the maximum thrust T of the turbojet engine at sea level and in a stationary state. n and its corresponding fuel consumption rate SFC n, converted to the fuel consumption rate at the cruise speed r (%), Mach number Ma, and altitude H (km); the changes in thrust and fuel consumption rate with flight speed, speed, and speed can be approximately expressed by the following fitting formula:

[0079]

[0080]

[0081] First, calculate the required flight thrust T based on the cruising speed and altitude. Then, interpolate the required thrust to obtain the required speed r (%). Finally, substitute the speed r (%) into the fuel consumption conversion formula to obtain the fuel consumption rate at the current altitude H, Mach number Ma, and thrust T.

[0082] (7) Based on the design indicators, solve the wing loading and thrust-to-weight ratio.

[0083] According to the minimum speed, static ceiling, maximum normal circling overload, maximum instantaneous overload and other indicators required in step (1), calculate the wing load for each working condition and take the maximum value as the wing load R W2S :

[0084]

[0085] Where n is the overload corresponding to the flight state (the minimum speed and static ceiling are the overloads in level flight state n=1, and the maximum normal circling overload and maximum instantaneous overload are the required values ​​in step (1)), V is the flight speed under the conditions of minimum speed, static ceiling, maximum normal circling overload and maximum instantaneous overload, ρ is the air density under the conditions of minimum speed, static ceiling, maximum normal circling overload and maximum instantaneous overload, and C ... L is the maximum available lift coefficient obtained in step (4).

[0086] According to the maximum flight speed, maximum climb angle and other indicators required in step (1), calculate the thrust-to-weight ratio of each working condition, and take the maximum value as the thrust-to-weight ratio of the cruise phase R T2W :

[0087]

[0088] C D is the drag coefficient under the conditions of maximum flight speed and maximum climb angle obtained in step (4), V is the maximum flight speed, θ ps is the maximum climb angle (0 in level flight), ρ is the air density of the corresponding working condition; the thrust-to-weight ratio obtained here is the local thrust-to-weight ratio and needs to be converted to the standard condition of zero altitude. Let H0 be the altitude under standard conditions, H be the current flight altitude of the aircraft, T wd0 is the standard altitude temperature (50°C is acceptable), then the temperature at altitude H is:

[0089] T wd =T wd0 -6.5 / 1000×(H-H0)

[0090] Among them, T wd and T wd0 The temperature unit is ℃; the static pressure at altitude H is:

[0091] p H =101325×(1-0.0065×H / 288.15) 5.2559

[0092] The speed of sound at altitude H is:

[0093]

[0094] Among them, the temperature T wd '=T wd The unit of +273.15 is K; then Ma=V / a, assuming γ=1.4, the total pressure p at altitude H is:

[0095]

[0096] The required thrust at altitude H0 (thrust-to-weight ratio R T2W0 )for:

[0097]

[0098] (8) Calculate the empty aircraft weight coefficient using the relative fuel mass and the relative fuel mass coefficient.

[0099] Empty weight factor R me The structural weight coefficient R me1 , booster engine case weight coefficient R me2 , fuel tank weight coefficient R me3 , turbojet engine weight coefficient R me4 The ratio of the structural weight to the total cruise weight is K. me1 (0.12~0.2), the ratio of booster fuel weight to total booster weight K me2 (0.68~0.8), fuel and tank weight ratio K me3 (0.68~0.85), turbojet engine single-machine thrust-to-weight ratio K me4 (3.7~10), and combined with the μ value calculated in step (5) and the R calculated in step (6) mfc , the empty weight coefficient can be calculated by the following formula:

[0100] R me1 =K me1 ×(1-μ)

[0101] R me2 =μ / K me2 -μ

[0102] R me3 =R mfc / K me3 -R mfc

[0103] R me4 =R T2W0 / K me4 ×(1-μ)

[0104] R me =R me1 +R me2 +R me3 +R me4

[0105] (9) Using the empty mass coefficient, fuel mass coefficient, and fuel mass coefficient in combination with design indicators, calculate the takeoff gross weight and the starting gross weight of the vertical launch vehicle in the cruise phase.

[0106] The empty mass coefficient R obtained from step (8) me , the booster fuel mass coefficient R obtained in step (5) mfv , the fuel mass coefficient R obtained in step (6) mfc and the payload mass M required in step (1) 1oad Calculate the aircraft's takeoff gross weight:

[0107]

[0108] The starting gross weight of the cruise segment is:

[0109] M1=M0×(1-μ)

[0110] (10) Using the design parameters, calculate the fuselage length of the vertical launch vehicle.

[0111] Assume the fuel tank length coefficient K L1 (take 1.73), boost length coefficient K L2 (take 1.56), turbojet engine body length coefficient K L3 (Take 12.1), and calculate the fuselage cross-sectional area S by combining the fuselage width / height (or fuselage diameter) determined in step (2) b , plus the instrument cabin length L determined in step (2) load , calculate the body length L b for:

[0112] L b =L load +(K L1 ×Rmfc ×M0 / S b +K L2 ×R mfv ×M0 / S b +K L3 ×R T2W0 ×M1) / 1000

[0113] (11) Determine the wing position and rudder wing size based on the design indicators and design parameters.

[0114] First, find the chord length c of the wing w for:

[0115]

[0116] Wingspan L w for:

[0117] L w =R l2c ×c w

[0118] Under subsonic conditions, the slope of the lift line of the airfoil can be determined by the semi-empirical formula:

[0119]

[0120] Among them, R l2c is the aspect ratio (when calculating the rudder wing, it is the rudder wing aspect ratio R l2c-wd ), η is the aerodynamic efficiency, which is taken as 0.9.

[0121] According to the static stability SM requirement of step (1), the slope of the torque curve can be calculated:

[0122] C mα =SM×C Lα

[0123] According to the control ratio Rc requirement in step (1), the control efficiency of the rudder surface can be calculated:

[0124]

[0125] The pitching moment generated by a unit rudder deflection angle can be calculated from the rudder surface lift generated by the unit rudder deflection angle and the distance from the rudder surface lift action point to the center of mass, and then the rudder surface lift generated by the unit rudder deflection angle can be calculated:

[0126]

[0127] Where x cg is the center of mass position (percentage of fuselage length), c w is the wing chord length.

[0128] According to the fuselage width b determined in step (2)ds , rudder wing aspect ratio R l2c-wd , using the same method as that used in step (11) to calculate C Lα The same formula is used to calculate the lift line slope C of the rudder wing Lα-wd , and then the rudder wing area can be calculated:

[0129]

[0130] The span and chord length of the rudder wing can be calculated from the rudder wing area and the rudder wing aspect-chord ratio.

[0131] The aerodynamic calculation method (such as AVL) in step (4) is used to calculate the static stability when the wing is installed at 40% to 60% of the fuselage, and a curve of static stability versus wing position is obtained. This curve is then used to interpolate and solve the wing position when the static stability is equal to the static stability index in step (1), as shown in the following example: Figure 3 shown.

[0132] (12) Iteratively calculate the shape parameters of the vertical launch vehicle until the iterative residual of each shape parameter is less than the preset requirement.

[0133] After the initial values ​​of the shape parameters are given, after calculations in steps (4) to (11), the values ​​of the new shape parameters after this iteration can be obtained: the fuselage length obtained in step (10), the wingspan, chord length, wing position, and rudder span length obtained in step (11), the takeoff gross weight and the starting gross weight of the cruise phase obtained in step (9), and the wing load obtained in step (7); the values ​​after the iteration are repeatedly brought into steps (4) to (11) for the next iterative calculation, and after obtaining the values ​​of each shape parameter after the next iteration, the values ​​are subtracted from the values ​​after this iteration to obtain the residuals of the two adjacent iterations of each shape parameter;

[0134] Repeat iterative steps (4) to (11) until the iterative residuals of all shape parameters are less than the specified value (determined based on experience, generally can be taken as 1 / 1000 of the convergence value. For example, if the calculated total weight is 100 kg, the residual requirement is 0.1 kg).

[0135] (13) Output the calculation results of the vertical launch vehicle's shape parameters and complete the overall parameter closed-loop design.

[0136] The output calculation results include:

[0137] The minimum drag coefficient, maximum lift coefficient, and maximum lift-to-drag ratio obtained in step (4);

[0138] The booster engine fuel mass coefficient obtained in step (5);

[0139] The fuel quality coefficient obtained in step (6);

[0140] Wing loading and thrust-to-weight ratio obtained in step (7);

[0141] The empty mass coefficient obtained in step (8);

[0142] The takeoff gross weight and the cruise phase starting gross weight obtained in step (9);

[0143] The fuselage length, booster engine body length, turbojet engine body length, and fuel tank length obtained in step (10);

[0144] The wingspan, chord length, wing position as a percentage of fuselage length, and rudder span / chord length obtained in step (11) are:

[0145] In summary, the present invention combines the design characteristics and differences of aerospace and aviation vehicles, effectively combines the vertical launch stage with the overall design of the vehicle, and forms a unified overall parameter design method for vertical launch vehicles; according to the indicator-type design input, the overall parameter design calculation results can meet the indicator requirements, and the various parameters are theoretically closed-loop and self-consistent; a booster engine design calculation model based on the relative quantity differential motion equation is created, and the vertical turning stage is effectively combined with the reconnaissance and patrol stage through wing loading, thrust-to-weight ratio, and turning trajectory parameters.

[0146] Example:

[0147] The overall parameter closed-loop design of a vertical launch vehicle is carried out using the design parameters and index requirements shown in Table 1 below:

[0148] Table 1 Design indicators and design parameters of a vertical launch vehicle (example)

[0149]

[0150]

[0151] In the population parameter solution example, three sets of initial parameters were set: a high deviation of 50%, a close deviation, and a low deviation of 50%, as shown in Table 2 below. This analysis analyzes the convergence characteristics of the population parameter estimation procedure when the initial values ​​deviate significantly. The values ​​in the table may change with each iteration. When the residual error of the initial parameters decreases to the given value within the iterative cycle, the iteration is considered to have converged.

[0152] Table 2 Iteration initial value (example)

[0153]

[0154] Under the above calculation conditions, six iterations were performed, and the results for the approximate values, high deviation, and low deviation are shown in Tables 3, 4, and 5. When the initial values ​​were close, geometric parameters such as wing loading, span, chord, wing position, and fuselage length converged in just one iteration, and gross weight converged in just two iterations. When the initial values ​​were 50% lower or higher, geometric parameters such as wing loading, span, and wing position converged to four significant figures in just three iterations, while launch stage gross weight, main stage gross weight, wing chord, and fuselage length converged to four significant figures in just two to four iterations.

[0155] Table 3 Calculation results of initial values ​​of approximate values

[0156]

[0157]

[0158] Table 4 Calculation results of initial high bias values

[0159]

[0160] Table 5 Calculation results of initial value of low bias

[0161]

[0162] The iterative results for different initial values ​​confirm that the overall parameter design calculation method established in this chapter can stably converge to a certain value under different initial values, and the calculation results are not affected by the initial values. The iterative convergence rate is very fast. When the deviation from the convergence value is 50%, only four iterations are required to converge the residual error to within 0.1%, and six iterations can converge the residual error to within 10⁻⁶. Other parameters in the design results are shown in Table 6.

[0163] Table 6 Overall parameter design results

[0164]

[0165] This overall parameter design method can calculate detailed overall parameters, external dimension parameters and main subsystem parameters such as booster engine, turbojet engine and fuel.

[0166] The above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit them. Although the present application has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. These modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the various embodiments of the present application, and should all be included in the scope of protection of the present application.

Claims

1. A closed-loop design method for overall parameters of a vertical launch vehicle, characterized in that: include: Determine the design indicators of the vertical launch vehicle and give the design parameters of the vertical launch vehicle; Set the initial values ​​of the vertical launch vehicle's shape parameters for iterative calculation and start the iterative calculation: Determine aerodynamic characteristic parameters based on the parametric shape of the vertical launch vehicle; Combined with the design parameters, the relative mass of the fuel of the vertical launch vehicle is determined by solving the differential equations of relative motion, and the fuel mass coefficient of the booster is determined; Calculating the relative fuel mass coefficient of the turbojet engine of the vertical launch vehicle based on the design indicators and design parameters; Based on the design indicators, solve the wing loading and thrust-to-weight ratio; Calculate the empty aircraft weight coefficient through the relative mass of fuel and the relative mass coefficient of oil; Calculate the takeoff gross weight and cruising phase starting gross weight of the vertical launch vehicle using the empty mass coefficient, fuel mass coefficient, and fuel mass coefficient in combination with design indicators. Using the design parameters, calculate the fuselage length of the vertical launch vehicle; Determine the wing position and rudder wing size based on the design indicators and design parameters; Repeat the above iterative calculation process for the shape parameters of the vertical launch vehicle until the iterative residual of each shape parameter is less than the preset requirement, and output the final shape parameters.

2. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Based on the parametric shape of the vertical launch vehicle, determine the aerodynamic characteristics, including: Based on the given wing installation angle / sweep angle / taper ratio, airfoil, and initial value of the shape parameters, the shape parameters are iteratively updated. The relevant dimensional parameters of the aircraft are described through formatting commands, and the grid is divided for calculation. After the calculation is completed, the fuselage drag model is used to correct the aerodynamic characteristic parameters including the lift coefficient and drag coefficient.

3. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Combined with the design parameters, the relative mass of the fuel of the vertical launch vehicle is determined by solving the relative mass motion differential equations, and the booster fuel mass coefficient is determined, including: The relative quantity motion differential equations are constructed by the relative quantity method and solved to obtain the curves of velocity and trajectory angle varying with relative mass coefficient. The interpolation method is used to intercept the relative mass coefficient when the trajectory inclination angle is less than the trajectory inclination angle at the end of the boost in the design parameters and the speed is greater than the speed at the end of the boost in the design parameters, and this relative mass coefficient is used as the booster fuel mass coefficient.

4. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Based on the design indicators and design parameters, the relative mass coefficient of fuel of the turbojet engine of the vertical launch vehicle is calculated, including: The flight profile of a vertical launch vehicle can be divided into the launch adjustment phase, the cruise phase, and the strike phase. Based on the mass of the vertical launch vehicle at launch, the mass after the boost ends, the mass when entering stable cruise, the mass during the terminal mission, and the mass at mission completion, an expression for the relative mass coefficient of fuel is constructed. The unknown quantity in the expression of the relative fuel mass coefficient is the ratio of the mass at mission completion to the mass at launch, which can be expressed as the product of the mass ratios of each mission segment. For the calculation of the mass ratio of the cruise segment, one state, either cruise or hover, is selected for design, or designs are performed for both states separately, and the design result with the larger relative fuel mass coefficient is taken. The mass ratio of the cruise segment is characterized with reference to the Breguet range-time formula. The fuel consumption rate used in the calculation is converted to the fuel consumption rate at the speed, Mach number, and altitude under cruise conditions based on the maximum thrust of the turbojet engine at sea level and in a stationary state and its corresponding fuel consumption rate.

5. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Based on the design indicators, solve the wing loading and thrust-to-weight ratio, including: Calculate the wing load for each operating condition based on the minimum speed, static ceiling, maximum normal circling overload, and maximum instantaneous overload required in the design indicators, and take the maximum value as the wing load; According to the maximum flight speed and maximum climb angle required in the design indicators, the thrust-to-weight ratio for each operating condition is calculated, and the maximum value is taken as the thrust-to-weight ratio for the cruise phase and converted to the standard condition of zero altitude.

6. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: The empty weight coefficient is calculated by the relative fuel mass and the relative fuel mass coefficient, including: The empty weight coefficient is calculated by combining the structural weight coefficient, the booster engine case weight coefficient, the fuel tank weight coefficient, and the turbojet engine weight coefficient, with the ratio of the structural weight to the cruise total weight, the ratio of the booster fuel weight to the booster total weight, the ratio of the fuel to the fuel tank total weight, the turbojet engine single-machine thrust-to-weight ratio, and the relative mass of the fuel and the relative mass coefficient of the fuel.

7. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Using the design parameters, calculate the fuselage length of the vertical launch vehicle, including: Based on the fuel tank length coefficient, booster length coefficient, and turbojet engine fuselage length coefficient, the fuselage cross-sectional area is calculated in combination with the fuselage width / height of the step design parameters, and the fuselage length is calculated by adding the instrument compartment length in the design parameters.

8. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: Combine the design indicators and design parameters to determine the wing position and rudder wing size, including: First, the chord length of the wing is calculated using the aircraft's takeoff gross weight, wing loading, and the wing aspect ratio in the design index, and the wingspan is determined. Secondly, the lift line slope of the airfoil is calculated using the aerodynamic efficiency, aspect ratio, and Mach number. The slope of the moment curve is determined by combining the lift line slope of the airfoil with the static stability in the design index. Combined with the control wall in the design index, the control efficiency of the rudder is calculated. Finally, the rudder lift is calculated using the rudder control efficiency, center of mass position, wing chord length, and fuselage length; the rudder area is determined using the rudder lift and the slope of the lift line of the rudder wing; the rudder span and chord length are calculated from the rudder area and the rudder aspect ratio.

9. The closed-loop design method for overall parameters of a vertical launch vehicle according to claim 1, characterized in that: After the initial values ​​of the shape parameters are given, a set of values ​​of the shape parameters after each iteration can be obtained, and the next iteration calculation is performed based on this. After the values ​​of each shape parameter after the next iteration are obtained, they are subtracted from the values ​​after the current iteration to obtain the residuals of the two adjacent iterations of each shape parameter; Repeat the iteration until the iterative residuals of all shape parameters are less than the specified value, completing the parameter closed-loop design.

10. A terminal device comprising a processor, a memory, and a computer program stored in the memory; characterized in that: When the processor executes the computer program, it implements the closed-loop design method for overall parameters of a vertical launch vehicle according to any one of claims 1 to 9.

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