CN111159942A - Method for calculating roll damping torque of winged aircraft based on steady simulation - Google Patents
Method for calculating roll damping torque of winged aircraft based on steady simulation Download PDFInfo
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Abstract
The invention provides a roll damping torque calculation method of an aircraft with wings based on steady simulation, which comprises the following steps: the method comprises the steps of firstly, determining the rotational symmetry parameters of the appearance of the aircraft, secondly, establishing a computational grid model containing periodic boundaries, thirdly, setting periodic boundary conditions, fourthly, calculating a flow field by using a rotating reference system method, and fifthly, obtaining the rolling moment M of the aircraftxSixthly, calculating the rolling damping moment coefficient C of the aircraftlp. The calculation method provided by the invention has the advantages of simple and convenient grid division and small calculation amount; by adopting the steady solving method, the defects of large computation amount and complicated process of the unsteady simulation 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 torque coefficient of the winged aircraft is calculated.
Description
Technical Field
The invention relates to the technical field of calculation of roll damping torque of an aircraft with wings, in particular to a roll damping torque coefficient calculation method based on steady numerical simulation at a zero attack angle.
Background
The roll damping characteristics of an aircraft are of great significance to the aerodynamic design of the aircraft. In the rolling process of the aircraft, the motion of the surface of the aircraft can drive the surrounding fluid to move, so that the circumferential shearing force appears on the surface of the aircraft, and the rolling motion of the aircraft is hindered. On the other hand, for an aircraft with wings, the rolling motion changes the local angle of attack of the wing, and thus changes the pressure distribution on the surface of the wing, so that the wing generates a blocking effect on the rolling. The blocking effect of the aerodynamic shape of the aircraft on the roll is an important parameter of the 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 torque of the aircraft need to be examined according to the requirement.
The method for acquiring the roll damping torque in the current engineering application mainly comprises two types: the method comprises a wind tunnel test method and a numerical simulation-based calculation simulation method.
The first type of wind tunnel test method is a relatively traditional test measurement means. The method obtains the aerodynamic characteristic parameters of the aircraft by blowing and measuring the force of the aerodynamic profile model of the aircraft. And the roll damping torque coefficient can be obtained by measuring the aerodynamic coefficient of the aircraft appearance model in a roll state. The wind tunnel test has high test cost and complex test flow, flow field details of the aircraft in a rolling state cannot be obtained, and the flow mechanism generated by the rolling damping torque is difficult to obtain.
The second type of numerical calculation means has the advantages of relatively low cost, capability of acquiring details of a flow field of the aircraft circumfluence and contribution to analysis of aerodynamic force generation reasons. However, for the winged aircraft in the rolling state, the streaming of the aircraft is in an unsteady state due to the continuous motion of the wings relative to the ground coordinate system. Therefore, for the outflow of the winged aircraft in the rolling state, an unsteady calculation method is often required for simulation. The unsteady calculation needs time propulsion, the calculated amount is large, parameters such as physical time step length and the like need to be determined, the steps are complicated, and the estimation efficiency of a numerical method on the rolling damping torque of the aircraft is limited.
Disclosure of Invention
The invention aims to provide an analysis method for steady calculation simulation of zero-attack-angle roll damping torque of a winged aircraft, and solves the problems of large calculated amount and complex process when Computational Fluid Dynamics (CFD) is used for estimating the roll damping torque of the winged aircraft. By using the technical scheme of the invention, the best simulation effect can be obtained on the roll damping torque of the external shape of the winged aircraft with the rotational symmetry at the zero attack angle, and the number of computational grids and the number of numerical iteration steps can be obviously reduced on the premise of not reducing the simulation precision. Through the mirror image calculation result, the technical scheme of the invention can obtain the flow field simulation result consistent with the steady calculation method, thereby providing a sufficient basis for flow field analysis.
In contrast, the invention provides a roll damping torque calculation method of an aircraft with wings based on steady simulation, which comprises the following steps: the method comprises the steps of firstly, determining the rotational symmetry parameters of the appearance of the aircraft, secondly, establishing a computational grid model containing periodic boundaries, thirdly, setting periodic boundary conditions, fourthly, calculating a flow field by using a rotating reference system method, and fifthly, obtaining the rolling moment M of the aircraftxSixthly, calculating the rolling damping moment coefficient C of the aircraftlp。
Wherein, the first step of determining the rotational symmetry parameter of the aircraft profile includes that the aerodynamic profile of the targeted winged aircraft is rotationally symmetric about the longitudinal axis of the aircraft, and if the aerodynamic profile of the aircraft is rotationally symmetric, the aircraft has a rotational symmetry angle
The value of (A) is obtained from the relationship of the geometrical shape of the aircraft, if the geometrical shape of the aircraft is rotated by an angle along its longitudinal axis
After the aircraft rotates, the shapes of the aircraft before and after the aircraft rotates are completely overlapped, then
For the rotational symmetry angle of the aircraft。
Wherein the second step of establishing a computational mesh model including periodic boundaries comprises determining rotational symmetry angles
Then according to
Dividing the aerodynamic computational grid of an aircraft by dividing the aircraft
A computational grid of regions and ensuring that boundaries of the computational domain along a normal to the aircraft surface satisfy a rotation angle of
Is rotationally symmetrical.
Setting periodic boundary conditions on two boundaries of the calculation domain along the surface normal direction of the aircraft, wherein the third step of setting the periodic boundary conditions comprises the step of setting the periodic boundary conditions on the two boundaries of the calculation domain along the surface normal direction of the aircraft, and the grids on the specified boundaries rotate relative to the longitudinal axis of the aircraft according to the rotation angle of the position relative to the longitudinal axis of the aircraft
Or
The flow field physics at the latter position is calculated versus the throughput.
The fourth step of calculating the flow field by using a rotating reference system method comprises the step of defining the rotating motion of the flow field by using the rotating reference system method; adding a non-inertial source term of a wall boundary motion parameter and a space flow field control equation influenced by the roll rate p to realize the steady solution of the rolling aircraft on the unsteady streaming; under a non-inertial reference frame fixedly connected to the aircraft, the aerodynamic force of the aircraft can be solved constantly.
The fifth step of obtaining the rolling moment M of the aircraftxIncluding, after convergence of the calculation, obtaining the roll force of the aircraft by integrating the pressure and viscous forces of the aircraft surfaceMoment MxWherein the calculation of the viscous force takes into account the velocity component of the motion of the object plane boundary itself.
Wherein, the sixth step is to calculate the roll damping moment coefficient C of the aircraftlpIncluding M if the aircraft profile is a plane-symmetric profile, i.e. p is 0x0, and a roll damping torque coefficient ClpThe roll moment M is calculated without changing with the roll ratexAnd ClpThe relationship of (1) is:
Mx＝qSDΩClp(1)
according to MxAnd ClpCan obtain ClpWherein q is dynamic pressure in the form of:
q＝1/2ρV2(2)
v is the incoming flow velocity, S is the reference area, D is the reference length, omega is the dimensionless roll rate, rho is the density:
according to the obtained ClpThe roll torque of the aircraft in the roll state can be given according to the roll rate.
The invention has the advantages of
1. The calculation method provided by the invention has the advantages of simple and convenient grid division and small calculation amount. By adopting the periodic boundary, the requirement can be met only by dividing the computational grid of a region which is a fraction of the complete aircraft shape for the aircraft shape meeting the rotational symmetry. The method can effectively reduce the complexity of roll damping torque value prediction and obviously reduce the calculated amount, thereby realizing the rapid prediction of the aerodynamic coefficient and improving the efficiency of the aerodynamic appearance design of the aircraft.
2. The invention adopts a steady solving method, and can avoid the defects of large computation amount and complicated process of unsteady simulation. The non-stationary calculation needs to carry out independence check on the physical time step length, and if a double-time propulsion method is adopted, the inner iteration number also needs to be independently verified. If the grid irrelevant inspection is carried out on the unsteady calculation problem, the influence of the grid on the calculation result is mutually coupled with the influence of the physical time step length, so that the verification process is complicated and fussy. The rotating reference system method 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 calculation amount of the steady solving is obviously smaller than that of the unsteady solving method based on time advance on the whole. Therefore, the scheme provided by the invention can better improve the estimation efficiency of the roll damping torque 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 roll damping torque coefficients under the same incoming flow condition and a plurality of different roll rates need to be calculated, the steady simulation method can adopt the calculation result of a certain roll rate as the initial values of other calculation states, so that the convergence process is accelerated, and the calculation efficiency is improved.
Detailed description of the preferred embodiment
The following description is made of specific embodiments of the present invention.
The invention discloses a roll damping torque calculation method of an aircraft with wings based on steady simulation, which comprises the following steps:
first step of determining the parameters of rotational symmetry of the profile of an aircraft
The solution of the invention requires that the aerodynamic profile of the aircraft must be characterized by rotational symmetry about the longitudinal axis of the aircraft, otherwise the solution of the invention cannot be used for calculation. If the aerodynamic profile of the aircraft is a rotationally symmetrical profile, it must have a rotationally symmetrical angle
The value of (A) can be obtained simply by the aircraft geometry relationship if the aircraft geometry is rotated by an angle along its longitudinal axis
After the aircraft rotates, the shapes of the aircraft before and after the aircraft rotates are completely overlapped, then
I.e. the angle of rotational symmetry of the aircraft.
Second step of establishing a computational mesh model including periodic boundaries
In determining the angle of rotational symmetry
Then can be based on
And dividing the aerodynamic computation grid of the aircraft. Wherein the calculation grid of the complete appearance of the aircraft is not required to be divided, only the division is required
A computational grid of regions. That is, the profile of the aircraft is formed by the fact that the profile of the aircraft is rotationally symmetrical about its longitudinal axis
The identical parts are arranged in a rotating way along the vertical axis, so that the computational grid of one of the areas only needs to be divided to meet the computational requirements of the invention. It is also because the technical solution of the present invention only directly calculates the complete aircraft
The overall computational complexity can be significantly reduced. In addition, in the mesh division process, the boundary of the calculation domain along the normal direction of the aircraft surface must be ensured to meet the rotation angle
Is rotationally symmetrical.
Third step defining cycle boundary conditions
Setting periodic boundary conditions on two boundaries of the calculation domain along the normal direction of the aircraft surface, namely specifying the grid on the boundary according to the rotation angle of the position relative to the longitudinal axis of the aircraft
(or
) The post-position flow field physics is calculated versus the throughput. The periodic boundary is used for establishing a connection between two boundaries of a calculation domain along the normal direction of the surface of the aircraft, and the essential principle is that for the rotationally symmetrical profile aircraft, the circumfluence is also about the longitudinal axis at a zero attack angle
The angle is rotationally symmetrical.
The fourth step is to calculate the flow field by using a rotating reference system method
The rotational motion of the flow field is defined by the rotating reference frame method. Namely, the influence of the roll rate p on the wall boundary motion parameters and the non-inertial source term of the space flow field control equation is added, so that the steady solution of the rolling aircraft on the unsteady streaming is realized. The essential reason for this is that although the winged aircraft is constantly moving relative to the ground in the roll state relative to the ground reference frame, the aircraft flow around exhibits unsteady flow. However, if the reference frame is fixed to the aircraft, when the angle of attack is zero, the aircraft flow around exhibits a steady state, so that the aerodynamic force of the aircraft can be solved for a steady state in a non-inertial reference frame (rotating reference frame) fixed to the aircraft.
The fifth step is to obtain the rolling moment M of the aircraftx
After the calculation converges, the roll moment M of the aircraft may be obtained by integrating the pressure and viscous forces of the aircraft surfacex. Wherein the calculation of the viscous force takes into account the velocity component of the motion of the object plane boundary itself.
Sixthly, calculating the roll damping moment coefficient C of the aircraftlp
If the aircraft profile is a plane-symmetrical profile, i.e. p is 0, then there is Mx0, and assuming roll damping torque coefficient ClpThe roll moment M is calculated without changing with the roll ratexAnd ClpThe relationship of (1) is:
Mx＝qSDΩClp(1)
according to MxAnd ClpCan obtain Clp. Wherein q is dynamic pressure in the form of:
q＝1/2ρV2(2)
v is the incoming flow velocity, S is the reference area, D is the reference length, omega is the dimensionless roll rate, rho is the density:
according to the obtained ClpThe roll torque of the aircraft in the roll state can be given according to the roll rate.
Examples
The specific implementation mode of the method for calculating the roll damping torque of the winged aircraft based on the steady simulation is as follows:
to calculate the profile (0.04572 for the aircraft diameter) of the American ANF (Army-Navy basic finner) four-wing aircraft at Mach 2.49 and Reynolds number of 1.86X 105Roll damping torque coefficient at zero angle of attack is an example.
First step of determining the parameters of rotational symmetry of the profile of an aircraft
The ANF is in the shape of a four-wing aircraft, the wings are symmetrical wing-shaped, and the aircraft body is in the shape of a rotation body, so that the ANF is in rotational symmetry along a longitudinal axis and rotates by a rotation angle
I.e. 90 degrees rotational symmetry.
Second step, establishing ANF shape mesh model containing period boundary
Because of the fact that
The calculation area to be divided being the complete aircraft profile
The roll damping torque coefficient of the complete aircraft can be obtained by only calculating one fourth of the external flow field of the aircraft. Wherein a quarter of the area mayTo select the area containing one full flap, or to select the area containing two half flaps and the aircraft body between the two half flaps. Because the grid division is complicated because the boundary at two sides of the calculation domain is intersected with the wing by selecting the calculation domain comprising half wings, the calculation domain comprising a whole wing is selected for grid division, wherein the wing is positioned in the middle of the calculation domain, and the position rotates +/-45 degrees along the longitudinal axis of the aircraft to reach the position of the boundary at two sides of the calculation domain. The boundary outside the calculation domain is a quarter cylinder, the boundaries on the two sides of the aircraft keep consistent in shape, and the boundaries are overlapped after being rotated by 90 degrees along the longitudinal axis of the aircraft. Because the supersonic speed problem is calculated, the aircraft is arranged in the front of a 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 grids on two sides of the aircraft are distributed consistently as much as possible.
Third step defining cycle boundary conditions
Because the aircraft is rotationally symmetric about the aircraft longitudinal axis by 90 degrees, the boundaries of the aircraft sides that are rotationally symmetric about the aircraft longitudinal axis by 90 degrees have the same flow field parameters, and the boundaries of the aircraft sides can be defined as periodic boundaries.
The fourth step is to calculate the flow field by using a rotating reference system method
The inflow and outflow boundaries are conventional far-field boundaries, and the object plane boundary defines the roll rate. The control equation considers the parameters of fluid compressibility, fluid viscosity, heat conduction rate and the like and is selected according to actual calculation requirements. And 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 rotation speed omega is 0.015, and the rotation speed p is 385.8 rad/s. And (5) performing steady solution, and iterating until the aerodynamic coefficient of the aircraft is stable to obtain a converged calculation result.
The fifth step is to obtain the rolling moment M of the aircraftx
Method for acquiring rolling moment M of aircraft by integrating pressure and viscous force of object planex：
Mx＝-0.0548349775N·m (1)
Since the aircraft is only four of the complete aircraftOne-half, the roll moment of the aircraft should be at M calculatedxMultiplying by 4 on the basis, then:
Mx＝-0.21933991N·m (2)
sixthly, calculating the roll damping moment coefficient C of the aircraftlp
Since the ANF shape is a plane-symmetric shape, M satisfies the condition that p is 0x0. According to the roll moment M of the aircraft as a wholexRoll damping moment coefficient C can be calculatedlp：
Clp＝Mx/(qSDΩ)＝-16.8682 (3)
Wherein the dynamic pressure q is:
q＝1/2ρV2＝11549.13Pa (4)
the reference area S is 0.001641732 (aircraft body cross-sectional area) and the reference length D is 0.04572.
Roll damping torque coefficient C obtained according to calculationlpThe roll moment coefficient of the ANF shape in a roll state under the attack angle of Mach 2.49 and zero degree can be estimated. And coefficient results can be obtained only by regularly calculating a quarter-sized area of the complete aircraft.
The invention discloses a roll damping torque calculation method 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 torque coefficient of the winged aircraft. The invention utilizes the geometrical characteristic that the aerodynamic shape of the winged aircraft meets the rotational symmetry about the longitudinal axis of the aircraft, and calculates the roll damping torque based on the periodic boundary, so that the size of the calculation domain and the number of the calculation grids are only a fraction of that of the complete aircraft; when the periodic boundary is adopted for calculation, the method defines and calculates the rolling motion of the aircraft by using a rotating reference system method, so that the calculation can be solved regularly, and further the calculation amount is reduced; the method for calculating the stable constant is adopted, when the working conditions needing to be calculated are more, the result of one working condition can be used as the initial value of other working conditions, and the calculation efficiency is further improved. The application of the periodic boundary calculation method in the invention includes, but is not limited to, four-wing aircraft, and the calculation of the roll damping moment coefficient of any aircraft profile (such as six-wing profile, eight-wing profile, empennage and canard combination rotationally symmetric about the longitudinal axis of the aircraft, and the like) satisfying the geometric profile and rotationally symmetric about the longitudinal axis of the aircraft is within the protection scope of the invention. In addition, the application of the calculation method in the invention includes, but is not limited to, the roll damping torque coefficient calculation of a plane-symmetric straight wing aircraft, and the roll damping torque coefficient calculation of a wing-provided aircraft profile (such as a rolling arc wing profile) which is not plane-symmetric but is rotationally symmetric about the longitudinal axis of the aircraft also belongs to the protection scope of the invention.
Claims (7)
1. A roll damping torque calculation method of a winged aircraft based on steady simulation is characterized by comprising the following steps:
firstly, determining the rotational symmetry parameters of the aircraft profile,
second, establishing a computational mesh model containing periodic boundaries,
thirdly, setting a period boundary condition,
fourthly, calculating the flow field by using a rotating reference system method,
fifthly, obtaining the rolling moment M of the aircraftx，
Sixthly, calculating a rolling damping moment coefficient C of the aircraftlp。
2. The method of claim, wherein the first step of determining a rotational symmetry parameter of the aircraft profile comprises,
the aerodynamic profile of the winged aircraft to which the method is directed is rotationally symmetrical about the longitudinal axis of the aircraft, and if the aerodynamic profile of the aircraft is rotationally symmetrical, the aircraft has an angle of rotational symmetry
The value of (A) is obtained by the aircraft geometry relationship if the aircraft geometry is along its longitudinal axisRotation angle
After the aircraft rotates, the shapes of the aircraft before and after the aircraft rotates are completely overlapped, then
Is the rotational symmetry angle of the aircraft.
3. The method of claim, wherein said second step of building a computational mesh model containing periodic boundaries comprises,
in determining the angle of rotational symmetry
Then according to
Dividing the aerodynamic computational grid of an aircraft by dividing the aircraft
A computational grid of regions and ensuring that boundaries of the computational domain along a normal to the aircraft surface satisfy a rotation angle of
Is rotationally symmetrical.
4. The method of, wherein the third step of setting a cycle boundary condition comprises,
setting periodic boundary conditions on two boundaries of the calculation domain along the normal direction of the aircraft surface, and specifying the grid on the boundary according to the rotation angle of the position relative to the longitudinal axis of the aircraft
Or
The flow field physics at the latter position is calculated versus the throughput.
5. The method of claim, wherein the fourth step of performing a flow field calculation using a rotating reference frame method comprises,
defining the rotation motion of the flow field by a rotation reference system method; adding a non-inertial source term of a wall boundary motion parameter and a space flow field control equation influenced by the roll rate p to realize the steady solution of the rolling aircraft on the unsteady streaming; under a non-inertial reference frame fixedly connected to the aircraft, the aerodynamic force of the aircraft can be solved constantly.
6. The method of claim wherein in the fifth step, obtaining an aircraft roll torque MxComprises the steps of (a) preparing a mixture of a plurality of raw materials,
after the calculation converges, the roll moment M of the aircraft is obtained by integrating the pressure and the viscous force of the aircraft surfacexWherein the calculation of the viscous force takes into account the velocity component of the motion of the object plane boundary itself.
7. The method of claim wherein in the sixth step, the aircraft roll damping torque coefficient C is calculatedlpComprises the steps of (a) preparing a mixture of a plurality of raw materials,
if the aircraft profile is a plane-symmetrical profile, i.e. p is 0, then there is Mx0, and a roll damping torque coefficient ClpThe roll moment M is calculated without changing with the roll ratexAnd ClpThe relationship of (1) is:
Mx＝qSDΩClp(1)
according to MxAnd ClpCan obtain ClpWherein q is dynamic pressure in the form of:
q＝1/2ρV2(2)
v is the incoming flow velocity, S is the reference area, D is the reference length, omega is the dimensionless roll rate, rho is the density:
according to the obtained ClpThe roll torque of the aircraft in the roll state can be given according to the roll rate.
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