CN111274648A - Distributed flight load design method for civil aircraft leading edge flap - Google Patents
Distributed flight load design method for civil aircraft leading edge flap Download PDFInfo
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- 238000012544 monitoring process Methods 0.000 claims abstract description 18
- 230000005489 elastic deformation Effects 0.000 claims abstract description 14
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- 238000009530 blood pressure measurement Methods 0.000 description 11
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Abstract
The application belongs to the technical field of aircraft flight load design, and particularly relates to a distributed flight load design method for a leading edge flap of a civil aircraft, which comprises the following steps: determining the calculation state of the severe flight load of the leading edge flap; performing full-aircraft flight load calculation without considering the influence of structural elastic deformation based on the nonlinear pneumatic pressure distribution data to obtain a full-aircraft flight load calculation result; setting a typical monitoring profile on the aircraft wing, and calculating structural deformation; reconstructing the three-dimensional shape of the airplane; performing fluid mechanics calculation on the deformed three-dimensional shape, and extracting a fluid mechanics calculation result of the leading edge flap part; and calculating to obtain the final distributed flight load of the leading edge flap of the civil aircraft. According to the method for designing the distributed flight load of the leading edge flap of the civil aircraft, the influence of the elastic deformation of the structure on the flight load is considered, the accuracy of designing the distributed flight load of the leading edge flap of the civil aircraft can be effectively improved, and the weight of the aircraft can be reduced.
Description
Technical Field
The application belongs to the technical field of airplane flight load design, and particularly relates to a distributed flight load design method for a leading edge flap of a civil airplane.
Background
The flight load design of civil aircraft is closely related to safety, reliability and economy. The flying load is designed accurately, the requirements on reliability and safety can be met, the structural weight can be reduced, and the maneuvering performance and the economy are improved.
When the civil aircraft leading-edge flap carries out flight load design, the traditional distributed flight load design method is that in the range of an aircraft flight envelope line, based on the use limitation of the aircraft and the leading-edge flap, nonlinear aerodynamic pressure distribution data (namely pressure measuring wind tunnel test data or computational fluid mechanics data) is used for carrying out flight load calculation. In order to ensure the absolute safety and reliability of the aircraft, the influence of the elastic deformation of the structure on the flight load is not considered in the calculation process, and the civil aircraft can generate large deformation after being loaded. Not only the leading edge flap of the airplane generates structural elastic deformation after being loaded, but also the leading edge flap generates additional deformation due to the deformation of the wing after being loaded. The traditional design method inevitably leads to the increase of the weight of the airplane structure, and reduces the maneuvering performance and the economy of the airplane. The invention provides a design method of distributed flight loads of a leading edge flap of a civil aircraft, which obtains more accurate distributed flight loads of the leading edge flap by considering the influence of elastic deformation of a structure on the flight loads.
Disclosure of Invention
In order to solve at least one of the above technical problems, the present application provides a method for designing distributed flight loads of a leading edge flap of a civil aircraft.
The application discloses a method for designing distributed flight loads of leading edge flaps of civil aircrafts, which comprises the following steps:
step one, determining a calculation state of the severe flight load of a leading edge flap;
secondly, performing full-aircraft flight load calculation without considering structural elastic deformation influence based on nonlinear pneumatic pressure distribution data to obtain a full-aircraft flight load calculation result;
setting a typical monitoring profile on the airplane wing, and calculating structural deformation;
reconstructing the three-dimensional shape of the airplane to obtain the deformed three-dimensional shape;
fifthly, performing fluid mechanics calculation on the deformed three-dimensional shape, and extracting a fluid mechanics calculation result of the leading edge flap part;
and step six, according to the severe flight load calculation state of the leading edge flap in the step one, multiplying the hydrodynamics calculation result of the leading edge flap part in the step five by the speed pressure to obtain the final distributed flight load of the leading edge flap of the civil aircraft.
According to at least one embodiment of the application, in the first step, the flight load change rule of the leading edge flap is combed according to the use limit of the airplane and the leading edge flap and the full airplane flight load calculation state, so that the severe flight load calculation state of the leading edge flap is determined.
According to at least one embodiment of the present application, in the second step, the method further includes:
and selecting nonlinear pneumatic pressure distribution data required by calculation according to the severe flight load calculation state.
According to at least one embodiment of the present application, the third step includes:
setting a typical monitoring profile on the airplane wing, performing static loading analysis based on a structure finite element model by using the calculation result of the full airplane flight load, and calculating the structural deformation so as to obtain the deformation condition of the monitoring profile.
According to at least one embodiment of the present application, the structural finite element model is constrained at the center of gravity, and the size and shape of the fuselage frame section, and the size and shape of each airfoil profile, remain unchanged under load.
According to at least one embodiment of the present application, in the fourth step, the three-dimensional shape of the airplane is reconstructed by using a least square method according to the deformation condition at the typical monitoring section.
According to at least one embodiment of the present application, in the fifth step, a computational fluid dynamics model is built according to the three-dimensional shapes before and after the deformation, and fluid dynamics calculation is performed.
According to at least one embodiment of the present application, the nonlinear aerodynamic pressure distribution data is manometric wind tunnel test data, wherein a grid closest to a position of a manometric point is selected from the computational fluid dynamics model to obtain an increment of computational fluid dynamics data before and after deformation.
According to at least one embodiment of the present application, between the step five and the step six, further comprising:
and correcting the fluid mechanics calculation result of the leading edge flap part in the flight envelope range of the airplane.
The application has at least the following beneficial technical effects:
according to the distributed flight load design method for the civil aircraft leading edge flap, the influence of the elastic deformation of the structure on the flight load is considered, so that the distributed flight load design accuracy of the civil aircraft leading edge flap is effectively improved, the weight of the aircraft is reduced, and a new idea is provided for the distributed flight load design of the civil aircraft leading edge flap.
Drawings
FIG. 1 is a flow chart of a distributed flight load design method for a leading edge flap of a civil aircraft according to the present application;
FIG. 2 is an exterior view of an aircraft according to an embodiment of the present application;
FIG. 3 is a schematic view of a monitoring profile arrangement of the present application;
FIG. 4 is a finite element model diagram of an aircraft structure according to the present application;
FIG. 5 is a schematic representation of a comparison of the three-dimensional shapes of the aircraft of the present application before and after deformation;
FIG. 6 is a diagram of a computational fluid dynamics model of the present application.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application and should not be construed as limiting the present application. All other embodiments that can be derived by a person skilled in the art from the embodiments given herein without making any creative effort shall fall within the protection scope of the present application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In order to realize the accuracy of the civil aircraft leading edge flap distributed flight load design, the application provides a correction principle of aerodynamic pressure distribution data before structure elastic deformation, and the corrected aerodynamic pressure distribution data is used for calculating the civil aircraft leading edge flap distributed flight load.
The method for designing the distributed flight load of the leading edge flap of the civil aircraft is further described in detail with reference to the attached drawings 1-6.
The application discloses a method for designing distributed flight loads of leading edge flaps of civil aircrafts, which comprises the following steps:
and step S101, determining the calculation state of the severe flight load of the leading edge flap.
Specifically, because structural strength design usually only needs to use severe flight loads, the calculation state of the severe flight loads of the leading edge flaps needs to be screened and determined according to the use limit of the airplane and the leading edge flaps, the calculation state of the full airplane flight loads is combined, and the change rule of the flight loads of the leading edge flaps is combed.
And S102, performing full-aircraft flight load calculation without considering the influence of structural elastic deformation based on the nonlinear aerodynamic pressure distribution data to obtain a full-aircraft flight load calculation result.
Specifically, according to the severe flight load calculation state, the nonlinear aerodynamic pressure distribution data required by calculation is selected, and the full-aircraft flight load calculation without considering the structural elastic deformation influence is carried out on the basis of the nonlinear aerodynamic pressure distribution data to obtain the full-aircraft flight load calculation result.
And S103, setting a typical monitoring profile on the airplane wing, and calculating structural deformation.
Specifically, a typical monitoring section is arranged on an airplane wing, the result of the calculation of the full airplane flight load is used, static loading analysis is carried out on the basis of a structure finite element model, and the structural deformation is calculated to obtain the deformation condition of the monitoring section. At the moment, the structural finite element model needs to be restrained at the gravity center, and the size and the shape of the section of the frame of the machine body and the size and the shape of each airfoil profile of the structural finite element model are ensured to be unchanged under the loaded condition.
And S104, reconstructing the three-dimensional shape of the airplane to obtain the deformed three-dimensional shape.
Specifically, according to the deformation condition at a typical monitoring section, the three-dimensional shape of the airplane is reconstructed by using a least square method, and the deformed three-dimensional shape is obtained.
And S105, performing fluid mechanics calculation on the deformed three-dimensional shape, and extracting a fluid mechanics calculation result of the front edge flap part.
Specifically, a computational fluid mechanics model is established according to the three-dimensional shapes before and after deformation, fluid mechanics calculation is carried out, and a fluid mechanics calculation result is obtained. Because the grid density of the computational fluid dynamics model is far greater than the pressure measurement point density of the pressure measurement wind tunnel test data, if the pneumatic pressure distribution data is the pressure measurement wind tunnel test data, the grid closest to the pressure measurement point position is selected (in the computational fluid dynamics model) to obtain the increment of the computational fluid dynamics data before and after deformation.
In addition, step S105 may be followed by:
and S1051, correcting the pneumatic pressure distribution data related to the severe flight load calculation state of the leading edge flap in the range of the flight envelope of the airplane. According to the past flight load design experience, if the flight load is calculated by directly using linear pneumatic pressure distribution data, the accuracy of the obtained flight load is not high compared with that obtained by using non-linear pneumatic pressure distribution data. Thus, the present application uses nonlinear aerodynamic pressure profile data when calculating severe leading edge flap flight loads. If the nonlinear aerodynamic pressure distribution data are computational fluid dynamics data, establishing a computational fluid dynamics model according to the deformed three-dimensional shape of the airplane, and directly using the obtained computational fluid dynamics data as corrected aerodynamic pressure distribution data; if the nonlinear pneumatic pressure distribution data is pressure measuring wind tunnel test data, subtracting the fluid mechanics calculation results before and after deformation to obtain the increment of the fluid mechanics calculation data before and after deformation, and superposing the increment on the pressure measuring wind tunnel test data to be used as the corrected pneumatic pressure distribution data.
And S106, according to the severe flight load calculation state of the leading edge flap in the step one, multiplying the hydrodynamics calculation result of the leading edge flap part in the step five by the speed pressure to obtain the final distributed flight load of the leading edge flap of the civil aircraft.
Specifically, the modified aerodynamic pressure distribution data may be multiplied by the velocity pressure to obtain the leading edge flap distributed flight load. Distributed flight loads can be divided into chordwise and spanwise distributions. According to the position characteristics of the leading edge flap on the airplane, if the pneumatic pressure distribution data is pressure measurement wind tunnel test data, the chord direction distribution load of the leading edge flap can be set as follows: on the pressure measurement section, the load of the front edge is the load of the pressure measurement point closest to the front edge, the load of the rear edge is the load of the pressure measurement point closest to the rear edge, and linear interpolation is carried out on the loads between the front edge and the pressure measurement point closest to the front edge, between the rear edge and the pressure measurement point closest to the rear edge and between the pressure measurement points; the spanwise distributed load of the leading edge flap can be set to: the distributed load of the pressure section nearest to the end face is taken as the distributed load of the leading edge flap side end face close to the fuselage, the distributed load of the pressure section nearest to the end face is taken as the distributed load of the leading edge flap side end face close to the wing tip, and the distributed loads between the leading edge flap side end face close to the fuselage and the pressure section nearest to the end face, between the leading edge flap side end face close to the wing tip and the pressure section nearest to the end face and between the pressure sections are subjected to chord equal proportion linear interpolation. If the aerodynamic pressure distribution data are computational fluid dynamics data, the chord-wise and span-wise distribution loads of the leading edge flap are directly set to be the result of multiplying the aerodynamic pressure distribution data by the velocity pressure because the grid density of the computational fluid dynamics model is large enough.
In conclusion, the aerodynamic pressure distribution data are corrected by considering the elastic deformation of the loaded actual structure, and the corrected aerodynamic pressure distribution data are used for calculating the distributed flight load of the leading edge flap of the airplane; the method not only effectively improves the accuracy of the civil aircraft leading edge flap distributed flight load design, is beneficial to weight reduction of the aircraft, but also provides a new idea for the civil aircraft leading edge flap distributed flight load design.
The method for designing distributed flight loads of leading-edge flaps of civil aircraft according to the invention is further described in the following by way of a specific example:
the example is a civil aircraft leading edge flap distributed flight load design, and the outline diagram of the aircraft is shown in figure 2. According to the civil aircraft leading edge flap distributed flight load design method provided by the invention, the civil aircraft leading edge flap distributed flight load design idea is as follows:
1) according to the use limitation of the airplane and the leading edge flap, the calculation state of the leading edge flap is combined with the calculation state of the full airplane flight load, the change rule of the leading edge flap flight load influenced by the elastic deformation of the structure is not considered in combing, the calculation state of the leading edge flap serious flight load is screened and determined, and the calculation state is shown in the table 1;
TABLE 1 leading edge flap Severe flight load calculation State
2) Selecting nonlinear pneumatic pressure distribution data required by calculation according to the calculation state table 1 in the step 1), and performing full-aircraft flight load calculation without considering the influence of structural elastic deformation based on the computational fluid mechanics data to obtain a full-aircraft flight load calculation result;
3) a typical monitoring profile is provided on an aircraft wing, and a schematic diagram of the monitoring profile layout is shown in FIG. 3. Using the calculation result of the full-aircraft flight load in the step 2), carrying out static loading analysis based on a structural finite element model (a structural finite element model diagram is shown in figure 4), and obtaining the deformation condition of the typical monitoring section of the aircraft under the loading;
4) reconstructing the three-dimensional shape of the airplane by using a least square method according to the deformation condition at the typical monitoring section, wherein the three-dimensional shape before deformation and the three-dimensional shape after deformation are opposite to each other, and the three-dimensional shape is shown in FIG. 5;
5) establishing a computational fluid mechanics model (as shown in fig. 6) according to the deformed three-dimensional shape, performing fluid mechanics computation to obtain a fluid mechanics computation result, and taking out a front edge flap part in the computation result;
6) and (3) multiplying the hydrodynamics calculation result of the leading edge flap part in the step 5) by the speed and pressure according to the calculation state table 1 in the step 1) to obtain the final distributed flight load of the leading edge flap of the civil aircraft.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (9)
1. A method for designing distributed flight loads of a leading edge flap of a civil aircraft is characterized by comprising the following steps:
step one, determining a calculation state of the severe flight load of a leading edge flap;
secondly, performing full-aircraft flight load calculation without considering the influence of structural elastic deformation based on the nonlinear pneumatic pressure distribution data to obtain a full-aircraft flight load calculation result;
setting a typical monitoring profile on the airplane wing, and calculating structural deformation;
reconstructing the three-dimensional shape of the airplane to obtain the deformed three-dimensional shape;
fifthly, performing fluid mechanics calculation on the deformed three-dimensional shape, and extracting a fluid mechanics calculation result of the leading edge flap part;
and step six, according to the severe flight load calculation state of the leading edge flap in the step one, multiplying the hydrodynamics calculation result of the leading edge flap part in the step five by the speed pressure to obtain the final civil aircraft leading edge flap distributed flight load.
2. The method for designing the distributed flight loads of the civil aircraft leading-edge flap of claim 1, wherein in the first step, the flight load change rule of the leading-edge flap is combed according to the use limit of the aircraft and the leading-edge flap and the calculation state of the full-aircraft flight load, so as to determine the calculation state of the severe flight load of the leading-edge flap.
3. The method for designing distributed flight loads for a civil aircraft leading edge flap according to claim 1, characterized in that in step two, it further comprises:
and selecting nonlinear pneumatic pressure distribution data required by calculation according to the severe flight load calculation state.
4. The method for designing distributed flight loads for a civil aircraft leading-edge flap according to claim 1, characterized in that it comprises, in step three:
setting a typical monitoring section on the airplane wing, performing static loading analysis based on a structure finite element model by using the calculation result of the full airplane flight load, and calculating the structural deformation so as to obtain the deformation condition of the monitoring section.
5. The method for designing distributed flight loads for a civil aircraft leading-edge flap according to claim 4, characterized in that the structural finite element model is constrained at the center of gravity, and the size and shape of the fuselage frame section and the size and shape of each airfoil profile remain unchanged under load.
6. The method for designing distributed flight loads of leading edge flaps of civil aircraft according to claim 1, characterized in that in step four, the three-dimensional profile of the aircraft is reconstructed using the least squares method according to the deformation at the typical monitoring profile.
7. The method for designing distributed flight loads of a civil aircraft leading-edge flap according to claim 1, characterized in that in the fifth step, a computational fluid dynamics model is established according to the three-dimensional shapes before and after deformation, and fluid dynamics calculation is carried out.
8. The method for designing the distributed flight loads of the civil aircraft leading-edge flap according to claim 7, wherein the nonlinear aerodynamic pressure distribution data are manometric wind tunnel test data, and wherein a grid closest to the manometric point in the computational fluid dynamics model is selected to obtain the increment of the computational fluid dynamics data before and after deformation.
9. The method for designing distributed flight loads for a civil aircraft leading edge flap according to claim 1, characterized in that it further comprises, between the fifth step and the sixth step:
and correcting the fluid mechanics calculation result of the leading edge flap part in the flight envelope range of the airplane.
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| CN112035947A (en) * | 2020-07-29 | 2020-12-04 | 成都飞机工业（集团）有限责任公司 | Method for calculating wing section load with integral oil tank |
| CN112214944A (en) * | 2020-10-27 | 2021-01-12 | 武汉理工大学 | Method for determining load of airplane subjected to wind shear caused by downburst during takeoff and landing |
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