CN111159942B - Method for calculating rolling damping moment of winged aircraft based on steady simulation - Google Patents
Method for calculating rolling damping moment of winged aircraft based on steady simulation Download PDFInfo
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Abstract
The invention provides a method for calculating rolling damping moment of a winged aircraft based on steady simulation, which comprises the following steps: the method comprises the steps of firstly, determining rotational symmetry parameters of the appearance of an aircraft, secondly, establishing a calculation grid model comprising a periodic boundary, thirdly, setting periodic boundary conditions, fourthly, calculating a flow field by using a rotational reference system method, and fifthly, obtaining the rolling moment M of the aircraft x Sixth step, calculating the roll damping moment coefficient C of the aircraft lp . The calculation method provided by the invention has the advantages of simple grid division and small calculation amount; by adopting the steady solving method, the defects of large unsteady analog calculation amount and complex process can be avoided. The steady simulation method provided by the invention has higher calculation efficiency when the calculation working conditions are more, and solves the problems of complex process and large calculation amount when the zero attack angle roll damping moment coefficient of the winged aircraft is calculated.
Description
Technical Field
The invention relates to the technical field of calculation of rolling damping moment of a winged aircraft, in particular to a rolling damping moment coefficient calculation method based on constant value simulation under zero attack angle.
Background
The roll damping characteristics of an aircraft are of great importance for the aerodynamic design of an aircraft. During the rolling process of the aircraft, the movement of the surface of the aircraft can drive the surrounding fluid to move, so that the surface of the aircraft generates circumferential shearing force, thereby preventing the rolling movement of the aircraft. On the other hand, for winged aircraft, the roll motion will change the local angle of attack of the wing, which in turn changes the pressure distribution on the wing surface, causing the wing to produce a roll retarding effect. The blocking effect of the aerodynamic shape of the aircraft on the roll is an important parameter of aerodynamic design of the aircraft, so that the roll damping coefficient of the aircraft needs to be obtained, and the reason and the mechanism of the roll damping moment of the aircraft are inspected according to the requirement.
The methods for obtaining the rolling damping moment in the current engineering application mainly comprise two types: firstly, a wind tunnel test method and secondly, a calculation simulation method based on numerical simulation.
The first wind tunnel test method is a more traditional test measurement means. The aerodynamic characteristic parameters of the aircraft are obtained by blowing and measuring the force of the aerodynamic appearance model of the aircraft. The roll damping moment coefficient can be obtained by measuring the aerodynamic coefficient of the aircraft appearance model in the roll state. The wind tunnel test has higher test cost and more complex test flow, and the flow field details of the aircraft in the rolling state cannot be obtained, so that the flow mechanism generated by the rolling damping moment is difficult to obtain.
The second type of numerical calculation means has the advantages of relatively low cost, capability of acquiring flow field details of the aircraft bypass flow and contribution to analysis of aerodynamic force generation reasons. However, in the case of a winged aircraft in the roll state, the airflow around the aircraft is in an unsteady state due to the continuous movement of the wings relative to the ground coordinate system. Therefore, for outflow of winged aircraft in roll state, it is often necessary to use unsteady calculation methods for simulation. Time pushing is needed for unsteady calculation, the calculated amount is large, parameters such as physical time step and the like are needed to be determined, steps are complicated, and the estimated efficiency of a numerical method on the rolling damping moment of the aircraft is limited.
Disclosure of Invention
The invention aims to provide an analysis method for performing steady calculation simulation on zero attack angle rolling damping moment of a winged aircraft, which solves the problems of large calculated amount and complex process when Calculating Fluid Dynamics (CFD) is used for estimating the rolling damping moment of the winged aircraft. By using the technical scheme of the invention, the rolling damping moment of the appearance of the wing-carrying aircraft with the appearance being rotationally symmetrical under the zero attack angle can obtain the best simulation effect, and the number of calculation grids and the number of numerical iteration steps can be obviously reduced on the premise of not reducing the simulation precision. By mirroring the calculation result, the technical scheme of the invention can obtain the flow field simulation result consistent with the unsteady calculation method, and provides a sufficient basis for flow field analysis.
In view of the above, the invention provides a method for calculating the roll damping moment of a winged aircraft based on steady simulation, which comprises the following steps: the method comprises the steps of firstly, determining rotational symmetry parameters of the appearance of an aircraft, secondly, establishing a calculation grid model comprising a periodic boundary, thirdly, setting periodic boundary conditions, fourthly, calculating a flow field by using a rotational reference system method, and fifthly, obtaining the rolling moment M of the aircraft x Sixth step, calculating the roll damping moment coefficient C of the aircraft lp 。
Wherein the first step of determining the rotational symmetry parameters of the aircraft profile comprises rotationally symmetric about the longitudinal axis of the aircraft with respect to the aerodynamic profile of the aircraft, and if the aerodynamic profile of the aircraft is rotationally symmetric, the aircraft has a rotational symmetry angle The value of (2) is obtained from the relation of the geometry of the aircraft, if the geometry of the aircraft is rotated by an angle of rotation +.>After that, the appearance of the aircraft before and after rotation is completely overlapped, then +.>Is the rotational symmetry angle of the aircraft.
Wherein the second step of establishing a computational grid model including periodic boundaries comprises, in determining rotational symmetry anglesAfter that, according to->Dividing aerodynamic calculation grid of aircraft, wherein the aircraft is required to be divided +.>A computational grid of regions and ensuring that the boundaries of the computational domain along the normal of the aircraft surface satisfy a rotation angle of +.>Is rotationally symmetrical.
Wherein the third step of setting the periodic boundary conditions comprises setting the periodic boundary conditions on two boundaries of the computational domain along the normal direction of the aircraft surface, the grid at the specified boundary being rotated by an angle relative to the longitudinal axis of the aircraft in terms of the positionOr (b)The physical flow field at the latter location calculates the convection flux.
The fourth step, the flow field calculation by using a rotary reference system method comprises the steps of defining the rotary motion of the flow field by the rotary reference system method; adding a non-inertial source item of a rolling speed p, which influences a wall boundary motion parameter and a space flow field control equation, and realizing steady solution of unsteady detour of the rolling aircraft; under a non-inertial reference frame fixedly connected to the aircraft, aerodynamic forces of the aircraft can be constantly solved.
Wherein, the fifth step is to obtain the rolling moment M of the aircraft x Comprising, after a convergence of the calculation, obtaining a roll moment M of the aircraft by integrating the pressure and the viscous force of the aircraft surface x Wherein the viscous force is calculated taking into account the motion velocity component of the object plane boundary itself.
Wherein, the sixth step is to calculate the roll damping moment coefficient C of the aircraft lp Including M if the aircraft shape is a plane symmetrical shape, i.e. p=0 x =0, and set a roll damping moment coefficient C lp Not changing with the rolling speed, the rolling moment M is calculated x And C lp The relation of (2) is:
M x ＝qSDΩC lp (1)
according to M x And C lp The relation of (C) can be obtained lp Wherein q is dynamic pressure in the form of:
q＝1/2ρV 2 (2)
v is the incoming flow speed, S is the reference area, D is the reference length, Ω is the dimensionless roll rate, ρ is the density:
according to the obtained C lp The roll moment of the aircraft in the rolled state can be given as a function of the roll rate.
The beneficial effects of the invention are that
1. The calculation method provided by the invention has the advantages of simple grid division and small calculation amount. With periodic boundaries, for aircraft shapes that meet rotational symmetry, the requirements can be met by dividing the computational grid by only a fraction of the area of the complete aircraft shape. The method can effectively reduce the complexity of the rolling damping moment numerical value estimation and obviously reduce the calculated amount, thereby realizing the rapid estimation of the pneumatic coefficient and improving the efficiency of the pneumatic appearance design of the aircraft.
2. The invention adopts a steady solving method, and can avoid the defects of large unsteady analog calculation amount and complex process. The unsteady calculation needs to carry out independence test on the physical time step, and if a double-time pushing method is adopted, the independent verification on the inner iteration number is also needed. If the grid independent test is performed on the unsteady calculation problem, the influence of the grid on the calculation result and the influence of the physical time step are mutually coupled, so that the verification process is complex and tedious. The rotating reference system legal steady solving scheme provided by the invention can effectively avoid the additional calculation process required by physical time step verification in unsteady calculation, and the calculated amount of steady solving is obviously smaller than that of the unsteady solving method based on time propulsion as a whole. Therefore, the scheme provided by the invention can better improve the estimated efficiency of the roll damping moment of the winged aircraft.
3. The steady simulation method provided by the invention has higher calculation efficiency when the calculation working conditions are more. When the rolling damping moment coefficients under the same incoming flow condition and a plurality of different rolling speeds need to be calculated, the steady simulation method can adopt the calculation result of a certain rolling speed as the initial value of the rest calculation states, so that the convergence process is accelerated, and the calculation efficiency is improved.
Description of the preferred embodiments
The following description is made of specific embodiments of the present invention.
The invention relates to a method for calculating the roll damping moment of a winged aircraft based on steady simulation, which comprises the following steps:
first step of determining rotational symmetry parameters of an aircraft shape
The solution according to the invention requires that the aerodynamic profile of the aircraft must have a rotationally symmetrical character with respect to the longitudinal axis of the aircraft, otherwise it is not possible to use the solution according to the invention for calculation. If the aerodynamic profile of an aircraft is rotationally symmetrical, it necessarily has a rotationally symmetrical angle The value of (2) can be obtained simply from the relationship of the aircraft geometry if the aircraft geometry is rotated by an angle of +.>After that, the appearance of the aircraft before and after rotation is completely overlapped, then +.>Namely the rotational symmetry angle of the aircraft.
Second step, establishing a calculation grid model containing periodic boundaries
In determining the angle of rotational symmetryAfter that, can be according to->The aerodynamic computational grid of the aircraft is partitioned. Wherein the calculation grid of the complete shape of the aircraft does not have to be divided, only the +.>A computational grid of regions. That is, since the shape of the aircraft is rotationally symmetrical about its longitudinal axis, the shape of the aircraft is defined by +.>The identical parts are arranged in a rotating way along the longitudinal axis, so that the computing grid of one area is only required to be divided, and the computing requirement of the invention can be met. It is also precisely because the solution according to the invention only directly calculates the complete aircraft +.>So that the overall calculation amount can be significantly reduced. Furthermore, during the meshing process, it must also be ensured that the boundary of the calculation domain along the normal of the aircraft surface satisfies a rotation angle of +.>Is rotationally symmetrical.
Third step of defining period boundary condition
Setting periodic boundary conditions on two boundaries of the computational domain along the normal direction of the aircraft surface, i.e. specifying the rotation angle of the grid at the boundary with respect to the longitudinal axis of the aircraft as a function of that position(or->) The flow field physical quantity at the rear position calculates the convection flux. The effect of the periodic boundary is to link the two boundaries of the computational domain along the normal to the aircraft surface, the essential principle being that for rotationally symmetrical profiled aircraft, its detouring is also +_ about the longitudinal axis at zero angle of attack>The angle is rotationally symmetrical.
Fourth step, flow field calculation is carried out by using a rotating reference system method
The rotational movement of the flow field is defined by a rotational reference frame method. The non-inertial source item affecting the wall boundary motion parameter and the space flow field control equation of the rolling speed p is added, and the steady solution of the unsteady detour of the rolling aircraft is realized. The essential reason for this is that although in roll state the wing of the winged aircraft is in constant motion relative to the ground with respect to the ground reference frame, the aircraft flow around therefore appears to be unsteady. However, if the reference frame is attached to the aircraft, the aircraft detour appears to be steady-state when the angle of attack is zero, so that the aerodynamic forces of the aircraft can be constantly solved in a non-inertial reference frame (rotating reference frame) attached to the aircraft.
Fifth step, obtaining the rolling moment M of the aircraft x
After the convergence of the calculation, the roll moment M of the aircraft can be obtained by integrating the pressure and the viscous force of the aircraft surface x . Wherein the viscous force is calculated taking into account the velocity component of the movement of the object plane boundary itself.
Sixth step, calculating the roll damping moment coefficient C of the aircraft lp
If the aircraft shape is a plane symmetrical shape, i.e. p=0 there is M x =0, and assuming a roll damping moment coefficient C lp Not changing with the rolling speed, the rolling moment M is calculated x And C lp The relation of (2) is:
M x ＝qSDΩC lp (1)
according to M x And C lp The relation of (C) can be obtained lp . Wherein q is dynamic pressure in the form of:
q＝1/2ρV 2 (2)
v is the incoming flow speed, S is the reference area, D is the reference length, Ω is the dimensionless roll rate, ρ is the density:
according to the obtained C lp The roll moment of the aircraft in the rolled state can be given as a function of the roll rate.
Examples
The specific implementation mode of the winged aircraft rolling damping moment calculation method based on the steady simulation is as follows:
to calculate the profile (0.04572 for aircraft diameter) of an ANF (Army-Navy basic finner) four wing aircraft in the United states at Mach 2.49 and Reynolds number 1.86×10 5 The roll damping moment coefficient at zero angle of attack is taken as an example.
First step of determining rotational symmetry parameters of an aircraft shape
The ANF shape is a four-wing aircraft shape, the wings adopt symmetrical wing sections, and the aircraft body is a rotating body shape, so that the ANF shape meets the rotation symmetry along a longitudinal axis and the rotation angleI.e. 90 degrees rotational symmetry.
Second step, building an ANF outline grid model containing periodic boundaries
Because ofThe calculation area to be divided is therefore +.>The roll damping moment coefficient of the complete aircraft can be obtained by only calculating one fourth of the aircraft outflow field. Wherein the quarter area may be selected to comprise a complete flap or may be selected to comprise two half flaps and an area of the aircraft body between the two half flaps. Since selecting a computing domain containing half-wings intersects boundaries on both sides of the computing domain with wings, resulting in a more complex meshing, selecting a computing domain containing a full-wing, where the wings are located, for meshingAnd calculating the middle position of the domain, wherein the position rotates by +/-45 degrees along the longitudinal axis of the aircraft, namely the position of the boundary on two sides of the calculated domain is reached. The outer boundary of the calculation domain is a quarter cylinder, the boundaries on two sides of the aircraft keep the same shape, and the boundaries are overlapped after rotating 90 degrees along the longitudinal axis of the aircraft. Because of the supersonic problem of calculation, the aircraft is placed in front of the calculation domain, the total length of the calculation domain is 15 times of the length of the aircraft, and the diameter of the calculation domain is 100 times of the diameter of the aircraft. And carrying out grid division on the basis of the calculation domain, and ensuring that the grid distribution on two sides of the aircraft is consistent as much as possible.
Third step of defining period boundary condition
Since the aircraft flow around is rotationally symmetric about the aircraft longitudinal axis by 90 degrees, the boundaries of the aircraft on both sides about the aircraft longitudinal axis by 90 degrees have consistent flow field parameters, and the boundaries on both sides of the aircraft can be defined as periodic boundaries.
Fourth step, flow field calculation is carried out by using a rotating reference system method
The inflow and outflow boundaries are conventional far field boundaries, and the object plane boundaries need to define roll-over rates. The control equation considers the compressibility of the fluid, and parameters such as the viscosity of the fluid, the heat conduction rate and the like are selected according to actual calculation requirements. And (3) solving a flow field control equation under a non-inertial system, wherein the rolling motion is around the longitudinal axis of the aircraft, the dimensionless rotating speed omega is 0.015, and the rotating speed p is 385.8rad/s. And (5) performing steady solving, and iterating until the aerodynamic coefficient of the aircraft is stable, so as to obtain a converged calculation result.
Fifth step, obtaining the rolling moment M of the aircraft x
Pressure and viscous force of integral object plane to obtain rolling moment M of aircraft x ：
M x ＝-0.0548349775N·m (1)
Since the aircraft is only one-fourth of the complete aircraft, the roll moment of the aircraft should be calculated as M x Multiplying 4 on the basis of the above, the following steps are:
M x ＝-0.21933991N·m (2)
sixth step, calculating the roll damping moment coefficient C of the aircraft lp
Due to the ANF profileIs a plane symmetrical shape which satisfies M when p=0 x =0. According to the roll moment M of the aircraft as a whole x Can calculate the roll damping moment coefficient C lp ：
C lp ＝M x /(qSDΩ)＝-16.8682 (3)
Wherein the dynamic pressure q is:
q＝1/2ρV 2 ＝11549.13Pa (4)
the reference area S is 0.001641732 (cross-sectional area of the aircraft body) and the reference length D is 0.04572.
According to the calculated rolling damping moment coefficient C lp The roll moment coefficient of the roll state of the ANF profile at Mach 2.49 and zero degree attack angle can be estimated. And coefficient results can be obtained by calculating the quarter-sized area of the complete aircraft at steady state.
The invention discloses a method for calculating roll damping moment of a winged aircraft based on steady simulation, which solves the problems of complex process and large calculation amount when calculating a zero attack angle roll damping moment coefficient of the winged aircraft. According to the invention, the aerodynamic shape of the winged aircraft is utilized to meet the geometric characteristics of rotational symmetry about the longitudinal axis of the aircraft, and the rolling damping moment is calculated based on a periodic boundary, so that the calculated domain size and the calculated grid number are only a fraction of those of the complete aircraft; the method uses a rotary reference system method to define and calculate the rolling motion of the aircraft while calculating by adopting a periodic boundary, so that the calculation can be solved steadily, and the calculated amount is further reduced; according to the steady computing method adopted by the invention, when the working conditions needing to be computed are more, the result of a certain working condition can be used as the initial value of other working conditions, so that the computing efficiency is further improved. The periodic boundary calculation method of the invention is applied to, but not limited to, four-wing aircrafts, and any calculation of the roll damping moment coefficient of the aircraft appearance (such as six-wing appearance, eight-wing appearance, tail wing and duck wing combination body rotationally symmetrical about the longitudinal axis of the aircraft, and the like) which meets the geometric appearance and is rotationally symmetrical about the longitudinal axis of the aircraft is within the protection scope of the invention. Furthermore, applications of the calculation method of the present invention include, but are not limited to, planar symmetric straight wing aircraft roll damping moment coefficient calculations, which are also within the scope of the present invention for non-planar symmetric winged aircraft profiles (e.g., cambered wing profiles) that are rotationally symmetric about the aircraft longitudinal axis.
Claims (1)
1. A method for calculating the roll damping moment of a winged aircraft based on steady simulation is characterized by comprising the following steps:
the first step, determining rotational symmetry parameters of the aircraft shape, comprising: the aerodynamic profile of the aircraft with wings aimed at is rotationally symmetrical about the longitudinal axis of the aircraft, the aircraft having a rotationally symmetrical angle if the aerodynamic profile of the aircraft is a rotationally symmetrical profile The value of (2) is obtained from the relation of the geometry of the aircraft, if the geometry of the aircraft is rotated by an angle of rotation +.>After that, the appearance of the aircraft before and after rotation is completely overlapped, then +.>Is the rotational symmetry angle of the aircraft;
the second step, establish the calculation grid model including periodic boundary, including: in determining the angle of rotational symmetryAfter that, according to->Dividing aerodynamic calculation grid of aircraft, wherein the aircraft is required to be divided +.>A computational grid of regions and ensuring that the boundaries of the computational domain along the normal of the aircraft surface satisfy a rotation angle of +.>Is rotationally symmetrical;
the third step, set up the periodic boundary condition, including: setting periodic boundary conditions on two boundaries of a computational domain along the normal direction of an aircraft surface, and designating a grid on the boundary according to the grid position to rotate an angle relative to the aircraft longitudinal axisOr- & lt- & gt>The physical flow field at the latter location calculates the convection flux.
Fourthly, calculating a flow field by using a rotation reference system method, wherein the method comprises the following steps; defining a rotational motion of the flow field by a rotational reference frame method; adding a non-inertial source item of a rolling speed p, which influences a wall boundary motion parameter and a space flow field control equation, and realizing steady solution of unsteady detour of the rolling aircraft; under a non-inertial reference system fixedly connected to the aircraft, aerodynamic force of the aircraft can be constantly solved;
fifthly, acquiring rolling moment M of aircraft x Comprising: after convergence of the calculation, the roll moment M of the aircraft is determined by integrating the pressure and the viscous force of the aircraft surface x Wherein the viscous force is calculated taking into account the motion velocity component of the object plane boundary itself.
Sixth step, calculating the roll damping moment coefficient C of the aircraft lp Comprising the steps of, in combination,
if the aircraft shape is a plane symmetrical shape, i.e. p=0 there is M x =0, and set a roll damping moment coefficient C lp Not changing with the rolling speed, the rolling moment M is calculated x And C lp The relation of (2) is:
M x ＝qSDΩC lp (1)
according to M x And C lp The relation of (C) can be obtained lp Wherein q is dynamic pressure in the form of:
q＝1/2ρV 2 (2)
v is the incoming flow speed, S is the reference area, D is the reference length, Ω is the dimensionless roll rate, ρ is the density:
according to the obtained C lp The roll moment of the aircraft in the rolled state can be given as a function of the roll rate.
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Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103968982A (en) * | 2014-05-09 | 2014-08-06 | 中国航天空气动力技术研究院 | Low damping rolling moment measuring device based on gas bearing |
| CN106354903A (en) * | 2016-08-18 | 2017-01-25 | 中国人民解放军国防科学技术大学 | Determination method of the computational domain outer boundary for solving the steady circumfluence value of flying objects |
| CN107451354A (en) * | 2017-07-27 | 2017-12-08 | 中国人民解放军军械工程学院 | A kind of emulation mode and terminal device of canard configuration rudders pneumatic power parameter |
| CN108121856A (en) * | 2017-12-06 | 2018-06-05 | 中国科学院力学研究所 | A kind of full flight domain aerocraft dynamic stability analysis method |
| CN108121855A (en) * | 2017-12-06 | 2018-06-05 | 北京理工大学 | Small-sized unmanned aircraft flight dynamics optimization method based on Bionic flexible wing |
| CN109063256A (en) * | 2017-06-30 | 2018-12-21 | 北京航空航天大学 | It is a kind of for assessing the airplane digital virtual flight simulation computing system of passenger plane airworthiness |
| CN110058603A (en) * | 2019-04-08 | 2019-07-26 | 北京电子工程总体研究所 | A kind of deformation reentry vehicle deformation is preceding to instruct torque to determine method |
| EP3518216A1 (en) * | 2018-01-30 | 2019-07-31 | Volvo Car Corporation | Co-simulation system with delay compensation and method for control of co-simulation system |
| CN110263497A (en) * | 2019-07-19 | 2019-09-20 | 南京航空航天大学 | A kind of pneumatic coupling influence analysis method based on relative gain |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8620492B2 (en) * | 2012-02-27 | 2013-12-31 | Textron Innovations Inc. | Yaw damping system and method for aircraft |
| US11132476B2 (en) * | 2017-12-13 | 2021-09-28 | Facebook, Inc. | Automatic aircraft design optimization based on joint aerodynamic, structural, and energy performance |
-
2019
- 2019-12-26 CN CN201911361153.8A patent/CN111159942B/en active Active
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103968982A (en) * | 2014-05-09 | 2014-08-06 | 中国航天空气动力技术研究院 | Low damping rolling moment measuring device based on gas bearing |
| CN106354903A (en) * | 2016-08-18 | 2017-01-25 | 中国人民解放军国防科学技术大学 | Determination method of the computational domain outer boundary for solving the steady circumfluence value of flying objects |
| CN109063256A (en) * | 2017-06-30 | 2018-12-21 | 北京航空航天大学 | It is a kind of for assessing the airplane digital virtual flight simulation computing system of passenger plane airworthiness |
| CN107451354A (en) * | 2017-07-27 | 2017-12-08 | 中国人民解放军军械工程学院 | A kind of emulation mode and terminal device of canard configuration rudders pneumatic power parameter |
| CN108121856A (en) * | 2017-12-06 | 2018-06-05 | 中国科学院力学研究所 | A kind of full flight domain aerocraft dynamic stability analysis method |
| CN108121855A (en) * | 2017-12-06 | 2018-06-05 | 北京理工大学 | Small-sized unmanned aircraft flight dynamics optimization method based on Bionic flexible wing |
| EP3518216A1 (en) * | 2018-01-30 | 2019-07-31 | Volvo Car Corporation | Co-simulation system with delay compensation and method for control of co-simulation system |
| CN110058603A (en) * | 2019-04-08 | 2019-07-26 | 北京电子工程总体研究所 | A kind of deformation reentry vehicle deformation is preceding to instruct torque to determine method |
| CN110263497A (en) * | 2019-07-19 | 2019-09-20 | 南京航空航天大学 | A kind of pneumatic coupling influence analysis method based on relative gain |
Non-Patent Citations (2)
| Title |
|---|
| 周强."考虑流固耦合效应的某飞行器力学性能分析".《应用力学学报》.2015,全文. * |
| 蒋胜矩."基于定常NS方程的飞行器滚转阻尼力矩系数导数计算方法".《弹 箭 与 制 导 学 报》.2008,全文. * |
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