Method for calculating the propulsion efficiency of an aircraft

Abstract

The application provides a kind of propulsive efficiency calculation method of aircraft, by obtaining the first target parameter of orbit transfer thruster, and determine the first ignition duration of orbit transfer thruster;Obtain the second target parameter of attitude control thruster, and determine the second ignition duration of attitude control thruster;According to the interference moment and control moment in the first target parameter and the second target parameter, determine the function relationship between the first ignition duration and the second ignition duration based on angular momentum conservation;According to the function relationship, the flow of orbit transfer thruster and the flow of attitude control thruster, calculate the consumption propellant proportion of the orbit transfer thruster and the attitude control thruster during the n time ignition;According to the consumption propellant proportion, calculate the thruster efficiency value during the n time ignition;According to the thruster efficiency value calculated during each ignition, obtain the total thruster efficiency value during total ignition.

Description

Technical Field

[0001] This invention relates to the field of aerospace technology, and in particular to a method for calculating the propulsion efficiency of an aircraft. Background Technology

[0002] The accuracy of satellite propellant budgets is crucial for ensuring the successful completion of satellite missions in orbit. Typically, satellites consume most of their propellant during orbit changes, therefore, it is necessary to study methods to improve the accuracy of propellant consumption budgets during satellite orbit changes.

[0003] Many factors can affect the accuracy of propellant budgets, among which orbital maneuvering efficiency is a crucial factor. However, the propulsion efficiency currently used in satellite design is mostly based on empirical values ​​derived from on-orbit statistical data, which may differ from the actual state of the satellite.

[0004] In summary, the existing methods have many problems in practical use, so it is necessary to improve them. Summary of the Invention

[0005] To address the aforementioned shortcomings, the present invention aims to provide a method and apparatus for calculating the propulsion efficiency of an aircraft, which can improve the accuracy of propellant budgeting, thereby facilitating the optimization of relevant parameters and improving the propulsion efficiency of the aircraft during trajectory changes.

[0006] To achieve the above objectives, the present invention provides a method for calculating the propulsion efficiency of an aircraft, comprising the following steps:

[0007] The first target parameters of the orbital changer thruster are obtained, and the first ignition duration of the orbital changer thruster is determined; wherein, the first target parameters include the disturbance torque and the flow rate of the orbital changer thruster generated during the nth ignition, and n is a positive integer;

[0008] The second target parameter of the attitude control thruster is obtained, and the second ignition duration of the attitude control thruster is determined; wherein, the second target parameter includes the control torque and the flow rate of the attitude control thruster generated during the nth ignition;

[0009] Based on the disturbance torque and the control torque, a functional relationship between the first ignition duration and the second ignition duration based on the conservation of angular momentum is determined;

[0010] Based on the functional relationship, the flow rate of the orbit changer thruster, and the flow rate of the attitude control thruster, calculate the propellant consumption ratio of the orbit changer thruster and the attitude control thruster during the nth ignition.

[0011] The thruster efficiency value during the nth ignition is calculated based on the propellant consumption ratio.

[0012] The total thrust efficiency value for the total ignition period is obtained based on the thrust efficiency values ​​calculated during each ignition period.

[0013] Optionally, the interference torque includes a first axial interference torque component and a second axial interference torque component, and the first axial interference torque component and the second axial interference torque component are distributed perpendicularly to each other along the axis; the control torque includes a first axial control torque component and a second axial control torque component, and the first axial control torque component and the second axial control torque component are distributed perpendicularly to each other along the axis; the second ignition duration includes a first axial ignition duration and a second axial ignition duration.

[0014] Optionally, let the first axial disturbance torque component be T. dxn The second axial disturbance torque component is T dyn The first ignition duration is Δt dn The first axial ignition duration is Δt cxn The second axial ignition duration is Δt cyn The second ignition duration Δt c =Δt cxn +Δt cyn The functional relationship is then determined based on the following formula:

[0015] |T dxn |Δt dn =|T cxn |Δt cxn ;

[0016] |T dyn |Δt dn =|T cyn |Δt cyn .

[0017] Optionally, the step of calculating the propellant consumption ratio of the orbital changer and the attitude control thruster during the nth ignition, based on the functional relationship, the flow rate of the orbital changer thruster, and the flow rate of the attitude control thruster, includes:

[0018] Based on the flow rate of the orbital changer thruster and the first ignition duration, calculate the first propellant consumption of the orbital changer thruster during the nth ignition.

[0019] The second propellant consumption of the attitude control thruster during the nth ignition is calculated based on the attitude control thruster flow rate and the second ignition duration.

[0020] Calculate the ratio of the first propellant consumption to the second propellant consumption to obtain the propellant consumption ratio of the orbital change thruster and the attitude control thruster during the nth ignition.

[0021] Optionally, let the first propellant consumption be ΔM. n The second propellant consumption is Δm n The flow rate of the variable-orbit thruster is I. fd The flow rate of the attitude control thruster is I. fc The propellant consumption ratio is p n The propellant consumption ratio is calculated based on the following formula:

[0022] ΔM n =KI fd Δt dn ;

[0023] Δm n =I fc (k1Δt cxn +k2Δt cyn );

[0024]

[0025] Wherein, K is the number of orbital change thrusters, k1 is the number of attitude control thrusters in the first axis, and k2 is the number of attitude control thrusters in the second axis.

[0026] Optionally, let the total propellant consumption during each ignition period of the orbital change thruster be ΔM, the total propellant consumption during each ignition period of the attitude control thruster be Δm, and the thruster efficiency during the nth ignition period be η. n The total thrust efficiency value η is calculated based on the following formula:

[0027] η n =1-p n ;

[0028]

[0029] Optionally, the first target parameter further includes the direction of the orbit-changing thrust required by the orbit-changing thruster and the first thrust of a single orbit-changing thruster; the second target parameter further includes the installation angle of the attitude control thruster and the second thrust of a single attitude control thruster.

[0030] After the step of calculating the propellant consumption ratio of the orbital changer and the attitude control thruster during the nth ignition based on the functional relationship, the flow rate of the orbital changer thruster, and the flow rate of the attitude control thruster, the method further includes:

[0031] Calculate the angle between the direction of the orbital change thrust and the installation angle;

[0032] Based on the included angle, the first thrust, and the second thrust, calculate the mass equivalent of the propellant consumed by the attitude control thruster for orbit change;

[0033] The propellant consumption ratio is corrected based on the mass equivalent value.

[0034] The step of calculating the thruster efficiency value during the nth ignition based on the propellant consumption ratio includes:

[0035] The thruster efficiency value during the nth ignition is calculated based on the corrected propellant consumption ratio.

[0036] Optionally, let the included angle be α, and the first thrust be F. dn The second thrust is F cn The mass equivalent value is Δm un The mass equivalent value is then calculated based on the following formula:

[0037]

[0038]

[0039] Among them, I fc For the attitude control thruster flow rate, I fd The flow rate of the variable-orbit thruster is Δm. n The amount of propellant consumed by the attitude control thruster during the nth ignition.

[0040] Optionally, let the modified propellant consumption ratio be P. ng The corrected propellant consumption ratio is calculated based on the following formula:

[0041]

[0042] Where, ΔM n The amount of propellant consumed by the orbital thruster during the nth ignition.

[0043] Optionally, let the modified thruster efficiency during the nth ignition be η. ng The corrected total thruster efficiency value η gThe corrected total thrust efficiency value is calculated based on the following formula:

[0044] η ng =1-p ng ;

[0045]

[0046] Where, Δt dn This is the first ignition duration.

[0047] This invention can obtain the theoretical propulsion efficiency based on the aircraft's control strategy and related propulsion parameters. Compared with the method of determining propulsion efficiency based on statistical laws, it improves the accuracy of efficiency estimation, thereby improving the accuracy of propellant budget. Furthermore, based on the calculated propulsion efficiency, it can propose inverse requirements for the design of relevant parameters affecting the propulsion efficiency of the aircraft during orbit changes, optimize the design of relevant parameters, and thus improve the propulsion efficiency of the aircraft during orbit changes. Attached Figure Description

[0048] Figure 1 A flowchart illustrating the steps of a method for calculating the propulsion efficiency of an aircraft according to an embodiment of the present invention. Detailed Implementation

[0049] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative and not intended to limit the invention.

[0050] It should be noted that references to "an embodiment," "embodiment," "example embodiment," etc., in this specification refer to the described embodiment including specific features, structures, or characteristics, but not every embodiment must include these specific features, structures, or characteristics. Furthermore, such expressions do not refer to the same embodiment. Moreover, when describing specific features, structures, or characteristics in conjunction with embodiments, whether or not explicitly described, it is indicated that incorporating such features, structures, or characteristics into other embodiments is within the knowledge of those skilled in the art.

[0051] Furthermore, certain terms are used in the specification and subsequent claims to refer to specific components or parts. Those skilled in the art will understand that manufacturers may use different names or terms to refer to the same component or part. This specification and subsequent claims do not distinguish components or parts by differences in name, but rather by differences in function. The terms "comprising" and "including" used throughout the specification and subsequent claims are open-ended and should be interpreted as "including but not limited to." Additionally, the term "connection" here includes any direct and indirect electrical connection means. Indirect electrical connection means include connections made through other means.

[0052] Figure 1 This invention illustrates a method for calculating the propulsion efficiency of an aircraft according to an embodiment of the present invention. The aircraft may be an artificial satellite, rocket, space probe, space shuttle, drone, etc.; this embodiment will be described below using an artificial satellite as an example. The method includes the following steps:

[0053] S101: Obtain the first target parameters of the orbit-changing thruster and determine the first ignition duration of the orbit-changing thruster; wherein, the first target parameters include the disturbance torque and flow rate of the orbit-changing thruster generated during the nth ignition, where n is a positive integer. The orbit-changing thruster is a device installed on an artificial satellite to perform satellite orbit-changing functions. The first target parameters can be obtained from the analysis of the performance and / or design working state of the orbit-changing thruster itself. For example, the flow rate and thrust of the orbit-changing thruster are determined by the thruster's own performance; the disturbance torque generated when the orbit-changing thruster ignites is determined by the thrust magnitude, the position of the satellite's center of mass, and the thruster's installation position; and the first ignition duration is derived from the analysis of the satellite's design working plan. The first ignition duration is the running time for the orbit-changing thruster to start ignition. The ignition duration specifically corresponds to the nth ignition period of the orbit-changing thruster. If the orbit-changing thruster has multiple ignition periods, it means that there are multiple corresponding first ignition durations.

[0054] S102: Obtain the second target parameters of the attitude control thruster and determine the second ignition duration of the attitude control thruster; wherein, the second target parameters include the control torque and flow rate of the attitude control thruster generated during the nth ignition. The orbit-changing thruster is a device installed on the artificial satellite to perform satellite attitude control. The second target parameters may be derived from the analysis of the performance and / or design of the attitude control thruster. For example, the control torque generated by the attitude control thruster is determined by the thrust magnitude, the position of the satellite's center of mass, and the thruster layout; the effect of the attitude control thruster on orbit change is determined by the angle between its direction and the direction of the orbit-changing thruster.

[0055] S103: Based on the disturbance torque and control torque, determine the functional relationship between the first ignition duration and the second ignition duration based on the conservation of angular momentum. That is, according to the angular momentum theorem, the relationship between the ignition duration of the attitude control thruster and the ignition duration of the trajectory change thruster can be determined.

[0056] Furthermore, the disturbance torque includes a first axial disturbance torque component and a second axial disturbance torque component, and the first axial disturbance torque component and the second axial disturbance torque component are distributed perpendicularly to each other along the axis; the control torque includes a first axial control torque component and a second axial control torque component, and the first axial control torque component and the second axial control torque component are distributed perpendicularly to each other along the axis; the second ignition duration includes a first axial ignition duration and a second axial ignition duration. Specifically, the first axial disturbance torque component corresponds to the X-axis in the coordinate system, and the second axial disturbance torque component corresponds to the Y-axis in the coordinate system; the first axial control torque component also corresponds to the X-axis in the coordinate system, and the second axial control torque component also corresponds to the Y-axis in the coordinate system.

[0057] This embodiment includes several attitude control thrusters, which are divided into X-axis attitude control thrusters and Y-axis attitude control thrusters according to the satellite design requirements. The first axial ignition duration specifically refers to the ignition duration of the X-axis attitude control thruster, and the second axial ignition duration refers to the ignition duration of the Y-axis attitude control thruster.

[0058] In one embodiment, let the first axial disturbance torque component be T. dxn The second axial disturbance torque component is T dyn The first ignition duration is Δt dn The first axial ignition duration is Δt cxn The second axial ignition duration is Δt cyn The second ignition duration Δt c =Δt cxn +Δt cyn The functional relationship is then determined based on the following formula:

[0059] |T dxn |Δt dn =|T cxn |Δt cxn ;

[0060] |T dyn |Δt dn =|T cyn |Δt cyn .

[0061] After determining the functional relationship between the first ignition duration and the second ignition duration based on the conservation of angular momentum, proceed to step S104.

[0062] S104: Based on the aforementioned functional relationship, the flow rate of the orbit changer thruster, and the flow rate of the attitude control thruster, calculate the propellant consumption ratio of the orbit changer thruster and the attitude control thruster during the nth ignition.

[0063] In one embodiment, step S104 includes:

[0064] Based on the flow rate of the orbit changer thruster and the first ignition duration, calculate the first propellant consumption of the orbit changer thruster during the nth ignition; based on the flow rate of the attitude control thruster and the second ignition duration, calculate the second propellant consumption of the attitude control thruster during the nth ignition; calculate the ratio of the first propellant consumption to the second propellant consumption to obtain the propellant consumption ratio of the orbit changer thruster and the attitude control thruster during the nth ignition.

[0065] Furthermore, let the first propellant consumption be ΔM. n The second propellant consumption is Δm n The flow rate of the variable-orbit thruster is I fd The attitude control thruster flow rate is I fc The propellant consumption ratio is p n The propellant consumption ratio is calculated based on the following formula:

[0066] ΔM n =KI fd Δt dn ;

[0067] Δm n =I fc (k1Δt cxn +k2Δt cyn );

[0068]

[0069] Where K represents the number of orbit-changing thrusters, k1 represents the number of attitude control thrusters in the first axis, and k2 represents the number of attitude control thrusters in the second axis. That is, k1 represents the number of X-axis attitude control thrusters, and k2 represents the number of Y-axis attitude control thrusters; the propellant consumption during the nth orbit-changing thruster ignition is ΔM. n The propellant consumption for attitude control during the nth thruster ignition is Δm. n .

[0070] S105: Calculate the thruster efficiency value during the nth ignition based on the propellant consumption ratio. This thruster efficiency value characterizes the apogee thruster efficiency of the aircraft during the nth ignition. Based on the above steps S101 to S105, the thruster efficiency values ​​corresponding to each ignition period can be calculated respectively.

[0071] S106: Obtain the total thrust efficiency value for the total ignition period based on the thruster efficiency values ​​calculated during each ignition period. Before the satellite enters its predetermined working orbit, the orbit-changing thruster typically ignites multiple times. Therefore, the total propellant consumption during the orbit-changing thruster ignition period is the sum of the propellant consumption during each ignition period, and the total attitude control propellant consumption during the ignition period is also the sum of the propellant consumption during each ignition period.

[0072] Specifically, let the total propellant consumption during each ignition period of the orbital change thruster be ΔM, the total propellant consumption during each ignition period of the attitude control thruster be Δm, and the thruster efficiency during the nth ignition period be η. n The total thrust efficiency value η is calculated based on the following formula:

[0073] η n =1-p n ;

[0074]

[0075] The total propellant consumption during each ignition period of the orbital change thruster is ΔM = ∑ n ΔM n The total propellant consumption during each ignition period of the attitude control thruster is Δm = ∑ n Δm n .

[0076] The total thrust efficiency value obtained in this embodiment represents the propulsion efficiency of the spacecraft, which can be regarded as providing input for the propulsion budget of the spacecraft. The method provided in this embodiment can be implemented in a pre-created logical model, that is, to establish a model for analyzing the propulsion efficiency of satellite orbit change, and to obtain the corresponding model output results by inputting the corresponding parameter information. Furthermore, the factors affecting the propulsion efficiency can be analyzed based on the model output results, so as to put forward counter-requirements for satellite design based on the influencing factors.

[0077] Considering that some satellites are designed with attitude control thrusters that also have orbit-changing capabilities, in such cases, a portion of the propellant consumed during attitude control will also contribute to orbit-changing. Therefore, when calculating the propellant consumed by attitude control, the portion used for orbit-changing needs to be subtracted from the propellant consumed by the attitude control thrusters.

[0078] To address this, an optional implementation is also provided, where the impulse generated by the attitude control thruster is equivalent to the impulse generated by the orbit change thruster, thus obtaining the portion of propellant consumed by the attitude control thruster that is equivalent to the orbit change portion; in this example, the first target parameter further includes the orbit change thrust direction required by the orbit change thruster and the first thrust of a single orbit change thruster; the second target parameter further includes the installation angle of the attitude control thruster and the second thrust of a single attitude control thruster; after step S104, the following is also included:

[0079] Calculate the angle between the direction of the orbit-changing thrust and the installation angle; based on the angle, the first thrust, and the second thrust, calculate the equivalent mass of the propellant consumed by the attitude control thruster for orbit changing; and correct the propellant consumption ratio based on the equivalent mass value.

[0080] Step S105 further includes: calculating the thruster efficiency value during the nth ignition based on the corrected propellant consumption ratio.

[0081] Specifically, let the included angle be α, and the first thrust be F. dn The second thrust is F cn The mass equivalent value is Δm un The mass equivalent value is then calculated based on the following formula:

[0082]

[0083]

[0084] Among them, I fc For attitude control of thruster flow, I fd For the flow rate of the variable-orbit thruster, Δm n The amount of propellant consumed by the attitude control thruster during the nth ignition.

[0085] Equivalent to the impulse generated by the attitude control thruster and the impulse generated by the trajectory change thruster, the following equivalent relationship can be obtained:

[0086]

[0087] Therefore, based on this equivalence relationship, the equivalent mass of the propellant consumed by the attitude control thruster for orbit change can be obtained; after obtaining the equivalent mass value, the propellant consumption ratio calculated in the aforementioned embodiment can be corrected.

[0088] Specifically, let the corrected propellant consumption ratio be P. ng The corrected propellant consumption ratio is calculated based on the following formula:

[0089]

[0090] Where, ΔM n The amount of propellant consumed by the orbital thruster during the nth ignition.

[0091] Furthermore, let the modified thruster efficiency during the nth ignition be η. ng The corrected total thruster efficiency value η g The corrected total thrust efficiency value is calculated based on the following formula:

[0092] η ng =1-p ng ;

[0093]

[0094] Where, Δt dn This is the first ignition duration.

[0095] Analysis revealed that the thruster's flow rate and thrust are determined by its own performance; the disturbance torque generated during the orbit-changing thruster's ignition is determined by the thrust magnitude, the satellite's center of mass position, and the thruster's installation location; the control torque generated by the attitude control thruster is determined by the thrust magnitude, the satellite's center of mass position, and the thruster's layout; and the attitude control thruster's effect on orbit changes is determined by its angle with the orbit-changing thruster's direction. By adjusting the satellite's center of mass and the thruster's installation location, angle, and number, propulsion efficiency can be improved to meet the satellite's requirements.

[0096] If the total thruster efficiency is lower than the preset requirement, the influencing factors of the orbital change propulsion efficiency can be analyzed according to the calculation formula, and then the propulsion efficiency can be improved by optimizing relevant parameters. Specifically, through analysis of the calculation results, the factors affecting the orbital change propulsion efficiency include: the flow rate, thrust, ignition duration and disturbance torque generated during ignition of the orbital change thruster, as well as the flow rate and control torque of the attitude control thruster.

[0097] Under normal circumstances, the flow rate and thrust of the thruster cannot be adjusted. However, by making corresponding requirements on the satellite's center of mass configuration and attitude control thruster layout, the interference torque of the orbit change thruster can be reduced, and the control torque of the attitude control thruster can be improved, so that the orbit change propulsion efficiency meets the satellite's needs.

[0098] This invention can analyze the mechanical relationships in the satellite orbit change process, and take into account different strategies for satellite orbit change and attitude control, as well as different situations such as the attitude control thruster also having orbit change functions, so as to analyze and obtain the calculation method of satellite single orbit change efficiency and total efficiency of multiple orbit changes.

[0099] Optionally, a propulsion efficiency calculation system for an aircraft is also provided, including a storage medium, a processor, and a computer program stored on the storage medium and executable on the processor, wherein the processor executes the computer program to implement any of the above-described propulsion efficiency calculation methods for the aircraft.

[0100] In summary, the propulsion efficiency calculation method for an aircraft described in this invention obtains the first target parameters of the orbit-changing thruster and determines the first ignition duration of the orbit-changing thruster; wherein, the first target parameters include the disturbance torque and the flow rate of the orbit-changing thruster generated during the nth ignition, and n is a positive integer; obtains the second target parameters of the attitude control thruster and determines the second ignition duration of the attitude control thruster; the second target parameters include the control torque and the flow rate of the attitude control thruster generated during the nth ignition; based on the disturbance torque and the control torque, a functional relationship based on the conservation of angular momentum is determined between the first ignition duration and the second ignition duration; based on the functional relationship, the flow rate of the orbit-changing thruster, and the flow rate of the attitude control thruster, the propellant consumption ratio of the orbit-changing thruster and the attitude control thruster during the nth ignition is calculated; based on the propellant consumption ratio, the thruster efficiency value during the nth ignition is calculated; based on the thruster efficiency values ​​calculated during each ignition period, the total thruster efficiency value for the total ignition period is obtained. Therefore, this invention can obtain the theoretical propulsion efficiency based on the control strategy and related propulsion parameters of the aircraft, which improves the accuracy of efficiency estimation compared to the method of determining propulsion efficiency based on statistical laws, thereby improving the accuracy of propellant budget; and based on the calculated propulsion efficiency, it can put forward reverse requirements on the design of relevant parameters affecting the propulsion efficiency of the aircraft during orbit change, optimize the design of relevant parameters, and thus improve the propulsion efficiency of the aircraft during orbit change.

[0101] Of course, the present invention may have other various embodiments. Without departing from the spirit and essence of the present invention, those skilled in the art can make various corresponding changes and modifications according to the present invention, but these corresponding changes and modifications should all fall within the protection scope of the appended claims.

Claims

1. A method for calculating the propulsion efficiency of an aircraft, characterized in that, Including the following steps: The first target parameters of the orbital changer thruster are obtained, and the first ignition duration of the orbital changer thruster is determined; wherein, the first target parameters include the disturbance torque and the flow rate of the orbital changer thruster generated during the nth ignition, and n is a positive integer; The second target parameter of the attitude control thruster is obtained, and the second ignition duration of the attitude control thruster is determined; wherein, the second target parameter includes the control torque and the flow rate of the attitude control thruster generated during the nth ignition; Based on the disturbance torque and the control torque, a functional relationship between the first ignition duration and the second ignition duration based on the conservation of angular momentum is determined; Based on the functional relationship, the flow rate of the orbit changer thruster, and the flow rate of the attitude control thruster, calculate the propellant consumption ratio of the orbit changer thruster and the attitude control thruster during the nth ignition. The thruster efficiency value during the nth ignition is calculated based on the propellant consumption ratio. The total thrust efficiency value for the total ignition period is obtained based on the thrust efficiency values ​​calculated during each ignition period.

2. The method for calculating the propulsion efficiency of an aircraft according to claim 1, characterized in that, The interference torque includes a first axial interference torque component and a second axial interference torque component, and the first axial interference torque component and the second axial interference torque component are distributed perpendicularly to each other along the axis; the control torque includes a first axial control torque component and a second axial control torque component, and the first axial control torque component and the second axial control torque component are distributed perpendicularly to each other along the axis; the second ignition duration includes a first axial ignition duration and a second axial ignition duration.

3. The method for calculating the propulsion efficiency of an aircraft according to claim 2, characterized in that, let... The first axial disturbance torque component is T dxn The second axial disturbance torque component is T dyn The first ignition duration is Δt dn The first axial ignition duration is Δt cxn The second axial ignition duration is Δt cyn The second ignition duration Δt c =Δt cxn +Δt cyn The functional relationship is then determined based on the following formula: |T dxn |Δt dn =|T cxn |Δt cxn ; |T dyn |Δt dn =|T cyn |Δt cyn 。 4. The method for calculating the propulsion efficiency of an aircraft according to claim 3, characterized in that, The step of calculating the propellant consumption ratio of the orbital changer and the attitude control thruster during the nth ignition, based on the functional relationship, the flow rate of the orbital changer thruster, and the flow rate of the attitude control thruster, includes: Based on the flow rate of the orbital changer thruster and the first ignition duration, calculate the first propellant consumption of the orbital changer thruster during the nth ignition. The second propellant consumption of the attitude control thruster during the nth ignition is calculated based on the attitude control thruster flow rate and the second ignition duration. Calculate the ratio of the first propellant consumption to the second propellant consumption to obtain the propellant consumption ratio of the orbital change thruster and the attitude control thruster during the nth ignition.

5. The method for calculating the propulsion efficiency of an aircraft according to claim 4, characterized in that, assuming... The first propellant consumption is ΔM n The second propellant consumption is Δm n The flow rate of the variable-orbit thruster is I. fd The flow rate of the attitude control thruster is I. fc The propellant consumption ratio is p n The propellant consumption ratio is calculated based on the following formula: ΔM n = KI fd Δt dn ; Δm n =I fc (k1Δt cxn +k2Δt cyn ); Wherein, K is the number of orbital change thrusters, k1 is the number of attitude control thrusters in the first axis, and k2 is the number of attitude control thrusters in the second axis.

6. The method for calculating the propulsion efficiency of an aircraft according to claim 5, characterized in that, assuming... The total propellant consumption during each ignition of the orbit-changing thruster is ΔM, the total propellant consumption during each ignition of the attitude control thruster is Δm, and the thruster efficiency during the nth ignition is η. n The total thrust efficiency value η is calculated based on the following formula: or n =1-p n ; 7. The method for calculating the propulsion efficiency of an aircraft according to claim 1, characterized in that, The first target parameter also includes the direction of the orbit-changing thrust required by the orbit-changing thruster and the first thrust of a single orbit-changing thruster; the second target parameter also includes the installation angle of the attitude control thruster and the second thrust of a single attitude control thruster. After the step of calculating the propellant consumption ratio of the orbital changer and the attitude control thruster during the nth ignition based on the functional relationship, the flow rate of the orbital changer thruster, and the flow rate of the attitude control thruster, the method further includes: Calculate the angle between the direction of the orbital change thrust and the installation angle; Based on the included angle, the first thrust, and the second thrust, calculate the mass equivalent of the propellant consumed by the attitude control thruster for orbit change; The propellant consumption ratio is corrected based on the mass equivalent value. The step of calculating the thruster efficiency value during the nth ignition based on the propellant consumption ratio includes: The thruster efficiency value during the nth ignition is calculated based on the corrected propellant consumption ratio.

8. The method for calculating the propulsion efficiency of an aircraft according to claim 7, characterized in that, assuming The included angle is α, and the first thrust is F. dn The second thrust is F cn The mass equivalent value is Δm un The mass equivalent value is then calculated based on the following formula: Among them, I fc For the attitude control thruster flow rate, I fd The flow rate of the variable-orbit thruster is Δm. n The amount of propellant consumed by the attitude control thruster during the nth ignition.

9. The method for calculating the propulsion efficiency of an aircraft according to claim 8, characterized in that, Let the corrected propellant consumption ratio be P. ng The corrected propellant consumption ratio is calculated based on the following formula: Where, ΔM n The amount of propellant consumed by the orbital thruster during the nth ignition.

10. The method for calculating the propulsion efficiency of an aircraft according to claim 9, characterized in that, set up The revised thrust efficiency value during the nth ignition is η. ng The corrected total thruster efficiency value η g The corrected total thrust efficiency value is calculated based on the following formula: or ng =1-p ng ; Where, Δt dn This is the first ignition duration.

Images