Launching scheme generation method, device and equipment of earth escape orbit and storage medium

Abstract

The invention discloses a launching scheme generation method, device and equipment of an earth escape orbit and a storage medium, and relates to the technical field of carrier rockets. The method comprises the following steps: determining injection parameters of detector injection, and determining an initial model selection of a carrier rocket according to the launching quality of a detector and the injection parameters; on the basis of the initial model selection, engine compartment parameters are determined according to the corresponding relation between the carrying capacity of the carrier rocket and C3, the maximum carrying capacity serves as the target, the launching flight path of the carrier rocket is determined according to the engine compartment parameters, and the carrier rocket carries a detection load; and generating a launch scheme of the carrier rocket according to the launch flight path and the engine compartment parameters. According to the technical scheme provided by the embodiment of the invention, the adaptability of the carrier rocket to launch the ultrahigh-energy escape orbit can be improved through optimal configuration of the power cabin, and a feasible launching scheme is provided for a deep space exploration task.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of launch vehicles, in particular to a method and device for generating a launch scheme of an earth escape orbit, equipment and a storage medium. BACKGROUND

[0002] Currently, deep space exploration missions usually adopt direct transfer or planet gravity assist transfer scheme. Considering the limitation of launch vehicle capacity, the C3 of the probe launched from the earth is generally less than 20 km 2 / s 2 However, for special tasks such as solar polar region exploration, a super high energy launch scheme needs to be adopted to achieve the fastest arrival near the exploration target.

[0003] Typically, the C3 of the probe launched from the earth by a super high energy launch scheme is usually more than 30 km 2 / s 2 For example, in the case of Neptune system exploration, a direct transfer probe achieves the orbit of Neptune, atmospheric entry and close-in exploration of Triton with a launch C3 of about 120 km 2 / s 2 , which can greatly shorten the arrival time and improve the mission reliability, and is of great significance to deep space exploration.

[0004] With the development of new generation manned lunar rocket and heavy launch vehicle and other more powerful launch vehicles, more choices are provided for super high energy launch of large mass deep space exploration payload. However, when the C3 is large, such as more than 100 km 2 / s 2 , the carrying capacity of the launch vehicle is poor when the launch vehicle carries the payload directly. SUMMARY

[0005] The present application provides a method and device for generating a launch scheme of an earth escape orbit, equipment and a storage medium to solve the problem of poor carrying capacity of the launch vehicle when the C3 is large.

[0006] In a first aspect, the present application provides a method for generating a launch scheme of an earth escape orbit, comprising:

[0007] determining an orbit insertion parameter of a probe, and determining an initial selection of a launch vehicle according to a launch mass of the probe and the orbit insertion parameter, wherein the C3 that can be reached by the launch vehicle carrying a payload is greater than a preset value;

[0008] determining a power cabin parameter according to the correspondence between the carrying capacity of the launch vehicle and the C3 based on the initial selection, and determining a launch flight trajectory of the launch vehicle according to the power cabin parameter with the maximum carrying capacity as the target, wherein the launch vehicle carries a detection payload.

[0009] generate a launch scheme of the launch vehicle according to the launch flight trajectory and the power cabin parameter.

[0010] In a second aspect, the present application provides a launch scheme generation device for an earth escape orbit, comprising:

[0011] a rocket selection module configured to determine an orbiting parameter of a probe, and determine an initial selection of a launch vehicle according to a launch mass of the probe and the orbiting parameter, wherein a C3 that can be reached by a launch load of the launch vehicle is greater than a preset value;

[0012] a flight trajectory determination module configured to determine a power cabin parameter according to a corresponding relationship between a carrying capacity and the C3 of the launch vehicle based on the initial selection, and determine a launch flight trajectory of the launch vehicle according to the power cabin parameter with a maximum carrying capacity as a target, wherein the launch vehicle carries a probe load;

[0013] a launch scheme generation module configured to generate a launch scheme of the launch vehicle according to the launch flight trajectory and the power cabin parameter.

[0014] In a third aspect, the present application provides an electronic device, comprising:

[0015] at least one processor;

[0016] and a memory in communication connection with the at least one processor;

[0017] wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor to enable the at least one processor to execute the launch scheme generation method for an earth escape orbit according to the first aspect.

[0018] In a fourth aspect, the present application provides a computer readable storage medium storing computer instructions, and the computer instructions are used to enable a processor to execute the launch scheme generation method for an earth escape orbit according to the first aspect.

[0019] The launch scheme generation method for an earth escape orbit provided by the present application selects a launch vehicle satisfying a C3 requirement according to a launch mass of a probe and an orbiting parameter, then determines a cabin configuration parameter based on a selection result of the launch vehicle, and then determines a launch flight trajectory according to the cabin configuration parameter with a maximum carrying capacity as a target, and finally generates a launch scheme of the launch vehicle with a load capacity according to the launch flight trajectory and the cabin configuration parameter. The launch scheme generation method can improve a carrying capacity of a launch vehicle launching an ultra-high energy orbit, and provide an excellent launch scheme for a deep space exploration mission.

[0020] It is to be understood that the embodiments described herein are merely exemplary of the application and that a person skilled in the art can devise other embodiments without departing from the scope of the present application. It is also to be understood that not all of the features and aspects of the present application are necessarily included in every embodiment of the application. In addition, the description of the embodiments of the present application is intended to cover any alternatives, modifications, equivalents and / or improvements of the embodiments of the present application which are within the spirit and scope of the present application as defined by the appended claims. BRIEF DESCRIPTION OF DRAWINGS

[0021] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings needed to be used in the embodiments will be briefly introduced. Obviously, the drawings in the following description are only some embodiments of the present application, and other drawings can be obtained by those skilled in the art without any creative effort based on these drawings.

[0022] Figure 1 is a flow chart of a launch scheme generation method of a geosynchronous escape orbit according to an embodiment of the present application;

[0023] Figure 2 is a flow chart of a launch scheme generation method of a geosynchronous escape orbit according to an embodiment of the present application;

[0024] Figure 3 is a structural schematic diagram of a launch scheme generation device of a geosynchronous escape orbit according to an embodiment of the present application;

[0025] Figure 4 is a structural schematic diagram of an electronic device according to an embodiment of the present application. DETAILED DESCRIPTION

[0026] In order to make the person skilled in the art better understand the present application, the technical solutions in the embodiments of the present application will be described clearly and completely with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only some embodiments of the present application, but not all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without any creative effort should be within the scope of the present application.

[0027] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application and the above-mentioned drawings are used to distinguish similar objects, and do not necessarily describe a specific order or sequence. It should be understood that the data thus used can be interchanged under appropriate circumstances, so that the embodiments of the application described herein can be implemented in an order other than that illustrated or described herein. In the description of the present application, "a plurality of" means two or more, unless otherwise specified. The association relationship of the associated objects is described, and it means that there can be three relationships, for example, A and / or B can mean that A exists alone, A and B exist together, and B exists alone. The character " / " generally represents an "or" relationship between the front and rear associated objects. In addition, the terms "include" and "have" and any variations thereof are intended to cover non-exclusive inclusion, for example, a process, method, system, product or device including a series of steps or units does not necessarily limit to those steps or units clearly listed, but can include other steps or units not clearly listed or inherent to these processes, methods, products or devices.

[0028] Embodiment one

[0029] Figure 1 A flowchart of a launch scheme generation method of a geocentric escape orbit is provided for the first embodiment of the present application. The present embodiment can be applied to the generation of a launch scheme of a carrier rocket. The method can be executed by a launch scheme generation device of a geocentric escape orbit, which can be realized in the form of hardware and / or software. The launch scheme generation device of a geocentric escape orbit can be configured in an electronic device, which can be composed of two or more physical entities or one physical entity.

[0030] As shown in Figure 1 The launch scheme generation method of a geocentric escape orbit provided by the first embodiment of the present application specifically includes the following steps:

[0031] S101, determining an orbit entry parameter of a probe, and determining an initial selection of a carrier rocket according to a launch mass of the probe and the orbit entry parameter, wherein the C3 value that can be reached by the carrier rocket launching a payload is greater than a preset value.

[0032] In the present embodiment, the orbit entry parameter of the probe launched from the earth can be determined according to the departure time before performing a deep space exploration task.

[0033] For example, first, the following Lambert problem is solved by iteration:

[0034]

[0035] Wherein, μ srepresents the solar gravitational constant, t2 and t1 represent the target time of arrival and the Earth departure time, respectively, and r1 and r2 represent the position vector of the Earth's center of mass in the heliocentric inertial coordinate system at the departure time and the position vector of the target celestial body's center of mass in the heliocentric inertial coordinate system at the arrival time, respectively. By solving the Lambert problem, the escape velocity increment of the Earth and the target capture velocity increment requirement at the target celestial body's sphere of influence at the arrival time can be obtained, and then the position and velocity of the escape point from the Earth's sphere of influence in the J2000 coordinate system can be obtained, which are denoted as r e and v e , respectively.

[0036] Taking r e and v e as initial values, the following flight dynamics equation is integrated:

[0037]

[0038] wherein r represents the vector from the Earth's center to the center of mass of the launch vehicle in the launch inertial system, represents the acceleration, μ S is the solar gravitational coefficient, μ j is the celestial body gravitational coefficient in the solar system, δ j is a switching function, F ns is a high-order perturbation term of gravitational perturbation, which is 1 when the probe is within the sphere of influence of the celestial body and 0 otherwise, F SRP represents the solar radiation pressure term, and F others represents other minor influence terms.

[0039] By fine-tuning the departure time velocity v e , the integral result of the above equation can be within the allowable range of the target point error. The specific correction method can be a differential correction algorithm. In the J2000 coordinate system, the corrected escape velocity and position of the Earth are denoted as r E and v E , respectively. Based on these parameters and the escape epoch, the orbit elements of the Earth's entry orbit, i.e., the entry parameters, can be obtained.

[0040] Then, according to the launch mass of the probe and the entry parameter requirement, the launch vehicle used for launch can be initially selected, i.e., a launch vehicle that can satisfy C3 greater than a preset value, such as a launch vehicle that can satisfy C3 greater than 30 km 2 / s 2 . C3 is the minimum velocity squared required to launch from the Earth's surface, overcome the Earth's gravity, and enter the target orbit.

[0041] Exemplarily, the way of determining the initial selection of the launch vehicle includes:

[0042]

[0043] where P, R and g are the thrust, aerodynamic force and gravity acceleration vectors in the inertial frame. and ψ(t) are the flight program pitch and yaw angles, which are the variables to be optimized. f f f p are the orbit inclination, the argument of perigee, the longitude of the ascending node and the perigee, respectively, which are the variables to be optimized. at the shutdown point. min-C3 indicates that the maximum C3 is targeted. The above formula indicates that, under the premise of determining the probe mass and the orbit injection parameters, the rocket with C3 greater than a preset threshold of 30 km / s is selected as the preliminary selected carrier rocket. 2 2

[0044] S102, based on the initial selection, determining the power cabin parameters according to the corresponding relationship between the carrying capacity of the carrier rocket and C3, and targeting the maximum carrying capacity, determining the launch flight trajectory of the carrier rocket according to the power cabin parameters, wherein the carrier rocket carries a detection load.

[0045] In this embodiment, the corresponding relationship between the carrying capacity of the carrier rocket of the initial selection and C3 can be determined first, and then the optimal power cabin parameters of the carrier rocket when carrying the detection load are determined according to the corresponding relationship, wherein the power cabin parameters can be understood as the parameters of the power cabin affecting the carrying capacity of the carrier rocket, such as the mass of the power cabin. Then, targeting the maximum carrying capacity, the launch flight trajectory of the carrier rocket is designed according to the power cabin parameters and the initial selection of the carrier rocket.

[0046] S103, generating the launch scheme of the carrier rocket according to the launch flight trajectory and the power cabin parameters.

[0047] In this embodiment, the launch scheme of the carrier rocket can be generated by using the launch flight trajectory and the power cabin parameters, so as to perform the deep space exploration task according to the scheme. The scheme can include the selection of the carrier rocket, the takeoff time, the taxiing time, the shutdown time, the launch flight trajectory and the power cabin parameters, etc.

[0048] ​​​​​The launch scheme generation method of the Earth escape orbit provided by the embodiment of the present application comprises the following steps: selecting a launch vehicle satisfying the C3 requirement according to the launch mass and the orbit injection parameter of the probe; determining the cabin configuration parameter based on the selection result of the launch vehicle; determining the launch flight trajectory of the launch vehicle according to the power cabin parameter, with the maximum launch capacity as the target; and finally generating the launch scheme of the launch vehicle with the load capacity according to the launch flight trajectory and the power cabin parameter.

[0049] Optionally, the launch flight trajectory of the launch vehicle is determined according to the power cabin parameter, with the maximum launch capacity as the target, which comprises the following steps: under the constraint of the dynamic equation of the launch vehicle, the launch flight trajectory of the launch vehicle is determined according to the power cabin parameter, with the maximum orbit injection mass corresponding to the launch capacity as the target.

[0050] For example, with the maximum launch capacity of the launch vehicle as the optimization target, the following optimization problem is solved based on the launch vehicle model and the power cabin parameter:

[0051]

[0052] C3 f =C3 exp ,i f =i E ,ω f =ω E ,Ω f =Ω E ,h p =h pE

[0053] Wherein, A0, t q , t h , t g and t f respectively represent the launch direction, the take-off time, the taxiing time and the shutdown time, C3 exp represents the escape C3 corresponding to the shutdown time, C3 E represents the escape C3 required for orbit injection, i E , ω E , Ω pE and h f represent the expected orbit inclination, the pericenter distance, the ascending node longitude and the perigee height.i f , ω f , Ω p and h f represent the orbit inclination, the pericenter distance, the ascending node longitude and the perigee at the shutdown time.m f represents the launch capacity function, min-mThis indicates that the optimization goal is to maximize the carrying capacity of the launch vehicle.

[0054] Example 2

[0055] Figure 2 This is a flowchart of a method for generating a launch scheme for an Earth escape trajectory, provided in Embodiment 2 of the present invention. The technical solution of the present invention is further optimized based on the above-mentioned optional technical solutions, and provides a specific method for generating a launch scheme for a carrier rocket.

[0056] Optionally, determining the propulsion module parameters based on the initial selection and the correspondence between the launch vehicle's carrying capacity and C3 includes: determining the trend of the initial selection launch vehicle's carrying capacity changing with C3, and determining the correspondence between the carrying capacity and C3 based on the trend, wherein the carrying capacity includes the ability to carry the propulsion module and the probe payload; under the constraint of the correspondence, determining the optimal propulsion module mass using the launch vehicle's acceleration capability, and determining propulsion module parameters other than the optimal propulsion module mass based on the optimal propulsion module mass, wherein the propulsion module parameters are propulsion module parameters that affect the carrying capacity.

[0057] Optionally, after generating the launch scheme of the launch vehicle based on the launch flight trajectory and the propulsion module parameters, the method further includes: performing Monte Carlo simulation on the launch scheme to obtain a first simulation result; if, based on the first simulation result, the lower bound of the 3δ boundary of the launch vehicle's orbital mass in the launch scheme is determined to be less than a preset expected payload capacity value, then adjusting the propulsion module parameters of the launch vehicle to obtain updated propulsion module parameters, wherein the lower bound is the difference between the mean of the orbital mass and three times the standard deviation; determining an updated launch scheme based on the updated propulsion module parameters, and performing Monte Carlo simulation on the updated launch scheme to obtain a second simulation result; if, based on the second simulation result, the lower bound of the 3δ boundary of the launch vehicle's orbital mass in the updated launch scheme is determined to be less than a preset expected payload capacity value, then adjusting the initial selection of the launch vehicle to obtain an updated initial selection; and determining the optimal launch scheme based on the updated initial selection.

[0058] like Figure 2 As shown in Embodiment 2 of the present invention, a method for generating a launch scheme for an Earth escape orbit specifically includes the following steps:

[0059] S201. Determine the orbital parameters for the probe to enter orbit, and determine the initial selection of the launch vehicle based on the launch mass of the probe and the orbital parameters.

[0060] S202, determine the change trend of the carrying capacity of the initial selected launch vehicle with C3, and determine the corresponding relationship between the carrying capacity and C3 according to the trend, wherein the carrying capacity includes the capacity of carrying the power cabin and the detection load.

[0061] Specifically, the change trend of C3 of the initial selected launch vehicle under different carrying capacities can be determined in advance, and the corresponding relationship between the carrying capacity and C3 can be determined according to the trend.

[0062] Optionally, the determination of the change trend of the carrying capacity of the initial selected launch vehicle with C3, and the determination of the corresponding relationship between the carrying capacity and C3 according to the trend, comprises: determining the change curve of the carrying capacity of the initial selected launch vehicle with C3, and performing curve fitting on the curve to obtain a fitting function of the carrying capacity and C3, wherein the fitting function represents the corresponding relationship between the carrying capacity and C3.

[0063] For example, under a given launch direction, the change curve of the carrying capacity of the initial selected launch vehicle with C3 under the condition that the power cabin is not driven can be determined first. Through curve fitting, a quadratic polynomial function of the carrying capacity and C3 can be obtained as follows:

[0064] M=a0+a1C3+a2C3 2

[0065] Wherein, M represents the carrying capacity, M=m load +m0, m0 is the mass of the power cabin in the launch vehicle, including the product mass of the support and control system, etc. load for the detection load.

[0066] S203, under the constraint of the corresponding relationship, determine the optimal power cabin mass by using the acceleration capacity of the launch vehicle, and determine the power cabin parameters other than the optimal power cabin mass according to the optimal power cabin mass, wherein the power cabin parameters are the power cabin parameters affecting the carrying capacity.

[0067] Specifically, the fitting function of the carrying capacity and C3 can be taken as a constraint condition, the maximum acceleration capacity of the initial selected launch vehicle is taken as a target, and the acceleration capacity function is solved to obtain the optimal power cabin mass. After determining the optimal power cabin mass, the corresponding power cabin parameters other than the optimal power cabin mass can be determined according to the mass.

[0068] Optionally, the determination of the optimal power cabin mass under the constraint of the corresponding relationship comprises: determining the sum of the perigee velocity of the launch vehicle at a preset height of the earth and the weighted value of the acceleration capacity under the constraint of the corresponding relationship, and determining the power cabin mass corresponding to the maximum sum as the optimal power cabin mass.

[0069] An exemplary structural coefficient (dry mass of the power cabin and the mass after refueling) is denoted as K, h p The characteristic velocity at the altitude can be expressed as:

[0070]

[0071] where V p is the perigee velocity, μ E is the Earth gravitational constant, and R0 is the average radius of the Earth. Similarly, the perigee velocity V pexp corresponding to the target orbit can be calculated. The velocity increment that the power cabin needs to achieve can be approximately expressed as ΔV = V pexp -V p . For the power cabin, if the propellant type is selected, the specific impulse range of the power cabin can be approximately determined, and the structural coefficient can be considered as a function of the propellant type and the size of the power cabin, i.e. is the propellant type, and i = 1, 2, 3, 4, 5 respectively represent liquid hydrogen and liquid oxygen, liquid oxygen and methane, liquid oxygen and kerosene, normal temperature propellant, and solid propellant. Under the premise of selecting the propellant, the main optimization parameter is the mass m0 of the power cabin. By solving the following single-parameter optimization problem, the optimal m0 can be obtained:

[0072]

[0073] s.t.m load +m0 = a0 + a1C3 + a2C3 2

[0074] where η is the loss coefficient. Considering the influence of the unusable amount of propellant in the power cabin and the safety margin, the consumption amount before the power cabin starts, and the velocity loss caused by the dynamics of the flight dynamics process, η can be taken as 0.85 to 0.9.I sp is the specific impulse, and g0 is the Earth gravitational acceleration. represents the acceleration capability.

[0075] Optionally, in addition to the optimal mass of the power cabin, the power cabin parameters at least include the propellant type and the specific impulse, and the mass of the power cabin includes the dry mass of the power cabin and the refueling mass.

[0076] S204, under the constraint of the dynamics equation of the launch vehicle, the maximum orbit entry mass corresponding to the carrying capacity is taken as the target, and the launch flight trajectory of the launch vehicle is determined according to the power cabin parameters.

[0077] S205, according to the launch flight trajectory and the power cabin parameters, a launch scheme of the launch vehicle is generated.

[0078] S206, performing Monte Carlo simulation on the launch scheme to obtain a first simulation result; if it is determined according to the first simulation result that a lower limit value of a 3δ boundary of an orbit injection mass of a carrier rocket in the launch scheme is less than a preset carrying capacity expectation value, adjusting a power cabin parameter of the carrier rocket to obtain an updated power cabin parameter.

[0079] The lower limit value is a difference between a mean value of the orbit injection mass and 3 times a standard deviation.

[0080] Specifically, if there are multiple initial selections of the carrier rocket, Monte Carlo simulation can be performed on a launch scheme corresponding to a currently selected carrier rocket. Specifically, the Monte Carlo simulation can be performed as follows:

[0081] According to the power cabin parameter and the launch flight trajectory in the launch scheme, and in combination with a deviation factor, Monte Carlo simulation is performed to determine whether the flight scheme is feasible, and a first simulation result is obtained. If the first simulation result shows that the lower limit value of the 3δ boundary of the orbit injection mass of the carrier rocket in the current launch scheme is greater than or equal to the preset carrying capacity expectation value, it indicates that the current launch scheme is feasible, and if the lower limit value is less than the preset carrying capacity expectation value, it indicates that the current launch scheme is infeasible. At this time, the power cabin parameter of the carrier rocket can be adjusted, such as changing the propellant type or adjusting the loss coefficient, to obtain an updated power cabin parameter.

[0082] The upper limit of the 3δ boundary is the mean value of the orbit injection mass + 3 times the standard deviation, and the lower limit is the mean value of the orbit injection mass - 3 times the standard deviation.

[0083] S207, determining an updated launch scheme based on the updated power cabin parameter, and performing Monte Carlo simulation on the updated launch scheme to obtain a second simulation result; if it is determined according to the second simulation result that a lower limit value of a 3δ boundary of an orbit injection mass of a carrier rocket in the updated launch scheme is less than a preset carrying capacity expectation value, adjusting an initial selection of the carrier rocket to obtain an updated initial selection.

[0084] Specifically, the launch scheme is regenerated using the updated power cabin parameter, and specific manners can be referred to the foregoing steps. Monte Carlo simulation is performed on the updated launch scheme to obtain a second simulation result. If the second simulation result shows that the lower limit value of the 3δ boundary of the orbit injection mass is still less than the preset carrying capacity expectation value, the initial selection of the carrier rocket can be adjusted, for example, if the initial selection includes a, b and c, and the current initial selection is a, the current initial selection can be adjusted to b to obtain an updated initial selection.

[0085] S208, determining an optimal launch scheme based on the updated initial selection.

[0086] Specifically, if the lower limit value of the 3delta boundary of the orbit injection quality of the updated initial selected type is still less than the preset carrying capacity expectation value, the initial selected type can be continuously adjusted until the initial selected type is traversed or the lower limit value is greater than or equal to the preset carrying capacity expectation value. When the lower limit value is greater than or equal to the preset carrying capacity expectation value, the current launch scheme is the optimal launch scheme.

[0087] The launch scheme generation method for the earth escape orbit provided by the embodiment of the application designs a launch scheme of a "carrier rocket + power cabin", increases the power cabin in the carrier rocket, accurately and reasonably determines the parameters of the power cabin, improves the adaptability of the carrier rocket to the super-high-energy deep space exploration launch task, effectively improves the carrying capacity of the rocket on the super-high-energy orbit, and verifies the rationality of the launch scheme through Monte Carlo simulation, thereby further ensuring the execution effect of the flight task.

[0088] Embodiment three

[0089] Figure 3 A structural schematic diagram of a launch scheme generation device for an earth escape orbit is provided in the embodiment three of the application. As shown in the figure, Figure 3 The device comprises a rocket selected type module 301, a flight trajectory determination module 302 and a launch scheme generation module 303, wherein:

[0090] The rocket selected type module is configured to determine an orbit injection parameter of a probe, and determine an initial selected type of a carrier rocket according to a launch mass of the probe and the orbit injection parameter, wherein the C3 that can be reached by the launch load of the carrier rocket is greater than a preset value.

[0091] The flight trajectory determination module is configured to determine a power cabin parameter according to a corresponding relationship between the carrying capacity and the C3 of the carrier rocket based on the initial selected type, and determine a launch flight trajectory of the carrier rocket according to the power cabin parameter with the maximum carrying capacity as the target, wherein the carrier rocket carries a detection load.

[0092] The launch scheme generation module is configured to generate a launch scheme of the carrier rocket according to the launch flight trajectory and the power cabin parameter.

[0093] The launch scheme generation device for the earth escape orbit provided by the embodiment of the application selects a carrier rocket that meets the C3 requirement according to the launch mass and the orbit injection parameter of the probe, then determines the cabin configuration parameter based on the selected type of the carrier rocket, determines the launch flight trajectory with the maximum carrying capacity as the target according to the power cabin parameter, and finally generates the launch scheme of the carrier rocket with the carrying capacity according to the launch flight trajectory and the power cabin parameter. The device can improve the carrying capacity of the carrier rocket on the super-high-energy orbit, and provides a feasible launch scheme for the deep space exploration task.

[0094] Optionally, the flight trajectory determination module comprises:

[0095] a corresponding relationship determination unit configured to determine a variation trend of the carrying capacity of the initial selected launch vehicle with respect to C3, and determine a corresponding relationship between the carrying capacity and C3 according to the trend, wherein the carrying capacity comprises the capacity of carrying the power cabin and the detection load.

[0096] a power cabin parameter determination unit configured to determine an optimal power cabin mass by using the acceleration capacity of the launch vehicle under the constraint of the corresponding relationship, and determine the power cabin parameters other than the optimal power cabin mass according to the optimal power cabin mass, wherein the power cabin parameters are the power cabin parameters affecting the carrying capacity.

[0097] Further, the determination of the variation trend of the carrying capacity of the initial selected launch vehicle with respect to C3, and the determination of the corresponding relationship between the carrying capacity and C3 according to the trend comprises: determination of a variation curve of the carrying capacity of the initial selected launch vehicle with respect to C3, and curve fitting of the curve to obtain a fitting function of the carrying capacity and C3, wherein the fitting function represents the corresponding relationship between the carrying capacity and C3.

[0098] Further, the power cabin parameters other than the optimal power cabin mass at least comprise propellant types and specific impulse, and the power cabin mass comprises dry mass of the power cabin and filling amount mass.

[0099] Further, the determination of the optimal power cabin mass by using the acceleration capacity of the launch vehicle under the constraint of the corresponding relationship comprises: determination of a sum value of a perigee velocity of the launch vehicle at a preset height of the Earth and a weighted value of the acceleration capacity under the constraint of the corresponding relationship, and determination of a power cabin mass corresponding to a maximum sum value as the optimal power cabin mass.

[0100] Optionally, the device further comprises:

[0101] a first simulation module configured to perform Monte Carlo simulation on the launch scheme to obtain a first simulation result after the generation of the launch scheme of the launch vehicle according to the launch flight trajectory and the power cabin parameters.

[0102] a first adjustment module configured to adjust the power cabin parameters of the launch vehicle to obtain updated power cabin parameters if a lower limit value of a 3δ boundary of an orbit injection mass of the launch vehicle in the launch scheme is less than a preset carrying capacity expectation value according to the first simulation result, wherein the lower limit value is a difference between a mean value and 3 times a standard deviation of the orbit injection mass.

[0103] The second simulation module is configured to determine an updated launch scheme based on the updated power cabin parameters, and perform Monte Carlo simulation on the updated launch scheme to obtain a second simulation result.

[0104] The second adjustment module is configured to adjust the initial selection of the launch vehicle to obtain an updated initial selection if a lower bound value of a 3delta boundary of an orbit injection mass of the launch vehicle in the updated launch scheme is less than a preset launch capacity expectation value according to the second simulation result.

[0105] The optimal launch scheme generation module is configured to determine an optimal launch scheme based on the updated initial selection.

[0106] Optionally, the flight trajectory determination module comprises:

[0107] The launch flight trajectory determination unit is configured to determine a launch flight trajectory of the launch vehicle according to the power cabin parameters under the constraint of a dynamic equation of the launch vehicle, with the maximum orbit injection mass corresponding to the launch capacity as the target.

[0108] The launch scheme generation device for the earth escape orbit provided by the embodiment of the application can execute the launch scheme generation method for the earth escape orbit provided by any embodiment of the application, and has the function modules and beneficial effects corresponding to the execution method.

[0109] Embodiment four

[0110] Figure 4 A structural schematic diagram of an electronic device 40 that can be used to implement embodiments of the application is shown. The electronic device is intended to represent various forms of digital computers, such as laptops, desktops, tablets, personal digital assistants, servers, blade servers, mainframes, and other appropriate computers. The electronic device can also represent various forms of mobile devices, such as personal digital assistants, cellular telephones, smart phones, wearable devices (e.g., headsets, glasses, watches, etc.), and other similar computing devices. The components shown here, their connections and relationships, and their functions, are meant to be examples only, and are not intended to limit the implementations of the applications described and / or claimed in this document.

[0111] As Figure 4As shown, the electronic device 40 includes at least one processor 41, and a memory, such as a read-only memory (ROM) 42, a random access memory (RAM) 43, and the like, connected to the at least one processor 41 in communication. The memory stores a computer program executable by the at least one processor 41, and the processor 41 can perform various appropriate actions and processes according to the computer program stored in the read-only memory (ROM) 42 or loaded into the random access memory (RAM) 43 from the storage unit 48. In the RAM 43, various programs and data required for the operation of the electronic device 40 can also be stored. The processor 41, the ROM 42, and the RAM 43 are connected to each other through a bus 44. An input / output (I / O) interface 45 is also connected to the bus 44.

[0112] Various components in the electronic device 40 are connected to the I / O interface 45, including an input unit 46, such as a keyboard, a mouse, and the like, an output unit 47, such as various types of displays, a speaker, and the like, a storage unit 48, such as a magnetic disk, an optical disk, and the like, and a communication unit 49, such as a network card, a modem, a wireless communication transceiver, and the like. The communication unit 49 allows the electronic device 40 to exchange information / data with other devices through a computer network, such as the Internet, and / or various telecommunication networks.

[0113] The processor 41 can be various general and / or special-purpose processing components with processing and computing capabilities. Some examples of the processor 41 include, but are not limited to, a central processing unit (CPU), a graphics processing unit (GPU), various special-purpose artificial intelligence (AI) computing chips, various processors running machine learning model algorithms, a digital signal processor (DSP), and any appropriate processor, controller, microcontroller, and the like. The processor 41 performs various methods and processes described above, such as the launch scenario generation method for a geocentric escape orbit.

[0114] In some embodiments, the launch scenario generation method for a geocentric escape orbit can be implemented as a computer program tangibly embodied in a computer readable storage medium, such as the storage unit 48. In some embodiments, part or all of the computer program can be loaded and / or installed onto the electronic device 40 via the ROM 42 and / or the communication unit 49. When the computer program is loaded into the RAM 43 and executed by the processor 41, one or more steps of the launch scenario generation method for a geocentric escape orbit described above can be performed. Alternatively, in other embodiments, the processor 41 can be configured to perform the launch scenario generation method for a geocentric escape orbit by any other appropriate means, such as by means of firmware.

[0115] The various embodiments of the systems and techniques described above can be implemented in digital electronic circuitry, integrated circuitry, a field programmable gate array (FPGA), an application specific integrated circuit (ASIC), a system on a chip (SOC), a complex programmable logic device (CPLD), computer hardware, firmware, software, and / or combinations thereof. These various embodiments can include implementation in one or more computer programs that are executable and / or interpretable on a programmable system including at least one programmable processor, which can be special or general purpose, coupled to receive data and instructions from, and to transmit data and instructions to, a storage system, at least one input device, and at least one output device.

[0116] Computer programs used to implement methods of the present application can be written in any combination of one or more programming languages. These computer programs can be provided to a processor of a general purpose computer, special purpose computer, or other programmable data processing apparatus to produce a machine, such that the computer program, when executed, implements the functions / acts specified in the flowcharts and / or block diagrams. The computer program can be executed entirely on a machine, partially on a machine, partially on a machine as part of a standalone software package, or entirely on a remote machine or server.

[0117] The computer device provided above can be used to execute the launch scheme generation method of the Earth escape orbit provided in any of the above embodiments, and has the corresponding functions and advantages.

[0118] Embodiment five

[0119] In the context of the present application, the computer-readable storage medium can be a tangible medium, the computer-executable instructions of which, when executed by a computer processor, are used to perform a launch scheme generation method of an Earth escape orbit, the method comprising:

[0120] determining an orbit injection parameter of a probe, and determining an initial selection of a launch vehicle according to a launch mass of the probe and the orbit injection parameter, wherein the initial selection corresponds to a C3 greater than a preset value;

[0121] based on the initial selection, determining a power cabin parameter according to a correspondence between a carrying capacity of the launch vehicle and C3, and determining a launch flight trajectory of the launch vehicle according to the power cabin parameter, with the goal of maximizing the carrying capacity;

[0122] generating a launch scheme of the launch vehicle according to the launch flight trajectory and the power cabin parameter.

[0123] In the context of the present application, a computer readable storage medium can be a tangible medium that can contain or store a computer program for use by or in connection with an instruction execution system, apparatus, or device. A computer readable storage medium can include, but is not limited to, an electronic, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any suitable combination of the foregoing. Alternatively, a computer readable storage medium can be a machine readable signal medium. More specific examples of a machine readable storage medium will include one or more lines of electrical connections, a portable computer diskette, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or Flash memory), an optical fiber, a portable compact disc read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the foregoing.

[0124] The computer device provided above can be used to execute the launch scheme generation method of the Earth escape orbit provided in any of the above embodiments, and has the corresponding functions and advantages.

[0125] It is worth noting that the above embodiments of the launch scheme generation device of the Earth escape orbit are only divided according to the functional logic, but are not limited to the above division, as long as the corresponding functions can be realized; in addition, the specific names of each functional unit are only for easy mutual differentiation, and do not limit the protection scope of the present application.

[0126] Note that the above are only preferred embodiments of the present application and the technical principles applied. Those skilled in the art will understand that the present application is not limited to the specific embodiments described herein, and those skilled in the art can make various obvious changes, readjustments and substitutions without departing from the scope of the present application. Therefore, although the present application has been described in more detail through the above embodiments, the present application is not limited to the above embodiments, and can include more other equivalent embodiments without departing from the concept of the present application, and the scope of the present application is determined by the scope of the appended claims.

Claims

1. A method for generating a launch scheme for an Earth escape trajectory, characterized in that, include: The orbital parameters for the probe to enter orbit are determined, and the initial selection of the launch vehicle is determined based on the launch mass of the probe and the orbital parameters, wherein the launch payload of the launch vehicle can achieve a C3 greater than a preset value. Based on the initial selection, the power module parameters are determined according to the correspondence between the launch vehicle's carrying capacity and C3. With the goal of maximizing carrying capacity, the launch trajectory of the launch vehicle is determined based on the power module parameters. The launch vehicle carries a probe payload. The launch plan for the launch vehicle is generated based on the launch flight trajectory and the power cabin parameters.

2. The method according to claim 1, characterized in that, Based on the initial selection, the determination of the power module parameters according to the correspondence between the launch vehicle's carrying capacity and C3 includes: The trend of the payload capacity of the initially selected launch vehicle with C3 is determined, and the correspondence between the payload capacity and C3 is determined based on the trend. The payload capacity includes the ability to carry the power module and the probe payload. Under the constraints of the aforementioned correspondence, the optimal power module mass is determined using the acceleration capability of the launch vehicle, and power module parameters other than the optimal power module mass are determined based on the optimal power module mass, wherein the power module parameters are those that affect the carrying capacity.

3. The method according to claim 2, characterized in that, The process of determining the trend of the payload capacity of the initially selected launch vehicle as a function of C3, and determining the correspondence between payload capacity and C3 based on the trend, includes: The curve of the payload capacity of the initially selected launch vehicle as a function of C3 is determined, and the curve is fitted to obtain the fitting function of payload capacity and C3, wherein the fitting function characterizes the correspondence between payload capacity and C3.

4. The method according to claim 2, characterized in that, The parameters of the power module other than the optimal power module mass include at least: propellant type and specific impulse, and the power module mass includes the dry weight mass and the propellant loading mass of the power module.

5. The method according to any one of claims 2-4, characterized in that, Determining the optimal power module mass using the acceleration capability of the launch vehicle, under the constraints of the aforementioned correspondence, includes: Under the constraints of the aforementioned correspondence, the weighted sum of the perigee velocity and acceleration capability of the launch vehicle at a preset altitude on Earth is determined, and the mass of the power cabin corresponding to the maximum sum is determined as the optimal power cabin mass.

6. The method according to claim 1, characterized in that, After generating the launch plan for the launch vehicle based on the launch flight trajectory and the power module parameters, the method further includes: A Monte Carlo simulation was performed on the launch scheme to obtain the first simulation result; If, based on the first simulation results, it is determined that the lower bound of the 3δ boundary of the orbital mass of the launch vehicle in the launch scheme is less than the preset expected value of the carrying capacity, then the power module parameters of the launch vehicle are adjusted to obtain updated power module parameters, wherein the lower bound is the difference between the mean of the orbital mass and 3 times the standard deviation. Based on the updated power pod parameters, an updated launch scheme is determined, and Monte Carlo simulation is performed on the updated launch scheme to obtain a second simulation result. If, based on the second simulation results, it is determined that the lower bound of the 3δ boundary of the orbital mass of the launch vehicle in the updated launch scheme is less than the preset expected carrying capacity, then the initial selection of the launch vehicle is adjusted to obtain the updated initial selection. The optimal launch scheme is determined based on the updated initial selection.

7. The method according to claim 1, characterized in that, The process of determining the launch trajectory of the launch vehicle based on the power module parameters, with the goal of maximizing carrying capacity, includes: Under the constraints of the launch vehicle's dynamic equations, with the goal of maximizing the orbital mass corresponding to the payload capacity, the launch trajectory of the launch vehicle is determined based on the parameters of the power module.

8. A device for generating a launch scheme for an Earth escape trajectory, characterized in that, include: The rocket selection module is used to determine the orbital parameters for the probe to enter orbit, and to determine the initial selection of the launch vehicle based on the launch mass of the probe and the orbital parameters, wherein the C3 that the launch vehicle's launch payload can achieve is greater than a preset value. The flight trajectory determination module is used to determine the power module parameters based on the initial selection and the correspondence between the launch vehicle's carrying capacity and C3, and to determine the launch flight trajectory of the launch vehicle based on the power module parameters with the goal of maximizing the carrying capacity, wherein the launch vehicle carries a detection payload; The launch scheme generation module is used to generate the launch scheme of the launch vehicle based on the launch flight trajectory and the power cabin parameters.

9. An electronic device, characterized in that, The electronic device includes: At least one processor; and A memory communicatively connected to the at least one processor; wherein, The memory stores a computer program executable by the at least one processor, which is executed by the at least one processor to enable the at least one processor to perform the launch scheme generation method for the Earth escape orbit as described in any one of claims 1-7.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer instructions that, when executed by a processor, implement the method for generating a launch scheme for an Earth escape orbit as described in any one of claims 1-7.

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