CN111929023B - Aircraft model driving system in wind tunnel and performance measuring method - Google Patents
Aircraft model driving system in wind tunnel and performance measuring method Download PDFInfo
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- CN111929023B CN111929023B CN202010802116.2A CN202010802116A CN111929023B CN 111929023 B CN111929023 B CN 111929023B CN 202010802116 A CN202010802116 A CN 202010802116A CN 111929023 B CN111929023 B CN 111929023B
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- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/08—Aerodynamic models
Abstract
The invention discloses a driving system and a performance measuring method for an aircraft model in a wind tunnel, belonging to the technical field of wind tunnel aerodynamics and flight dynamics experiments, wherein the driving system comprises a support rod device, a six-rod parallel mechanism and a linear guide rail, wherein the support rod device is positioned at the rear part of the aircraft model and supports the aircraft model through a multi-component balance in the model; a rolling driving and measuring device is arranged in the supporting rod device; the six-rod parallel mechanism comprises six connecting rods, two ends of each connecting rod are provided with universal hinges, and the universal hinges at the front ends are connected with the supporting rod devices; the universal hinge at the rear end is connected with the slide block of the linear guide rail; the linear guide rail is arranged on the inner wall of the wind tunnel experiment section. The invention can provide omnibearing, large-range and high-speed motion in the wind tunnel, has enough rigidity, strength and low interference, and can be used for measuring the aerodynamic force and virtual flight performance of an aircraft with high maneuverability and agility in the high-speed wind tunnel.
Description
Technical Field
The invention relates to the technical field of wind tunnel aerodynamics and wind tunnel flight dynamics (wind tunnel virtual flight) experiments, in particular to an aircraft model driving system in a wind tunnel and a method for measuring wind tunnel aerodynamics and wind tunnel virtual flight mechanical properties.
Background
High maneuverability, high agility and over-stall maneuverability are important indicators of modern high performance military aircraft and tactical missiles. To achieve this requirement, the aircraft should have excellent wide angle range of variation, angular velocity range of variation and instantaneous angular acceleration, as well as stability and control capability for over-stalled flight. In these complex maneuvers, the aircraft longitudinal and lateral three-dimensional flow disturbances, severe non-linearity and unsteady characteristics make the aerodynamic and flight dynamics of the aircraft very complex and difficult to find using computational aerodynamic or engineering computational methods. Therefore, the development of comprehensive, wide-range and high-speed dynamic test equipment in wind tunnel experiments is one of necessary prerequisites for obtaining aerodynamic data of an aircraft model.
Obtaining aerodynamic characteristics in maneuvering flight, particularly nonlinear, unsteady and cross-interfering characteristics, provides only initial parameters for aircraft design. In the conventional aircraft design method, aerodynamic data of the aircraft are obtained through wind tunnel experiments or computational fluid dynamics, so that the control system design and the overall aircraft design are carried out. However, during flight at large angles of attack and large maneuvers, the aerodynamic performance of the aircraft and the flight dynamics of the aircraft are no longer separately manageable due to unsteady, nonlinear and cross-interference in the aerodynamic and flight dynamics of the aircraft. The aircraft produced by the traditional aircraft design method may not only be far from the actual performance, but also be dangerous in flight tests.
Therefore, in order to develop an aircraft with high maneuverability and agility, besides improving the test capability of wind tunnel aerodynamics, a wind tunnel flight dynamics test or a wind tunnel virtual flight test must be developed. That is, as the wind tunnel test, if it can be classified into a low-level stage of the wind tunnel test (wind tunnel aerodynamic test) and a high-level stage of the wind tunnel test (wind tunnel flight dynamic test or wind tunnel virtual flight test), the wind tunnel test must be able to perform both of these tests.
The low-level stage of the wind tunnel test (wind tunnel aerodynamic test) is to develop comprehensive, large-range and high-speed dynamic test equipment so as to obtain wind tunnel test aerodynamic data of the aircraft model.
The advanced stage of the wind tunnel test (wind tunnel flight dynamics test or wind tunnel virtual flight test) is a comprehensive test of the flight dynamics characteristics, the flight control system and the navigation system of the aircraft in the wind tunnel. The device and the test method for comprehensively processing the aerodynamics, the flight dynamics and the control system are important means for providing the flight quality of the aircraft, reducing the risk of a flight test and shortening the development period of the aircraft. The wind tunnel virtual flight test technology can be used for providing aerodynamic force data of the aircraft, directly providing flight dynamics data of the uncontrolled aircraft, and verifying the accuracy, efficiency and reliability of hardware and control program software of a control system of the controlled aircraft under flight conditions.
In order to realize a complete wind tunnel flight dynamics test (wind tunnel virtual flight test), several requirements must be met: (a) the model has complete freedom degrees, namely the freedom degrees of three angles and three linear displacement directions; (b) the range of motion in these several degrees of freedom must be large enough to encompass the range of motion of an aircraft in maneuvering flight; (c) more importantly, the movement speed of the mechanism must be fast enough to reflect the movement speed of the aircraft in real time; (d) the corresponding flight dynamics analysis must simulate the dynamics data such as mass distribution, inertia and the like of a real aircraft; (e) the control surface and the control speed of the model should meet the speed ratio requirement of a real aircraft; (f) in order to carry out the test of the high-speed wind tunnel, the supporting rigidity of the model must be large enough; (g) for the versatility of the test, the test can be used for different types of tests (e.g. different types of aircraft or missiles, etc.) with little change in the mechanics. In short, to realize a "forced" wind tunnel virtual flight test in a wind tunnel, it is required that corresponding test equipment must be capable of providing all-directional freedom of motion; a sufficiently large variation range of the motion parameter; in particular, the controllable movement speed of the test device must be as large as possible to meet the control requirements of the flight dynamics. That is, the test apparatus must have all-round, wide-range, and high-rate performance, which is rarely provided by the prior art.
Disclosure of Invention
The invention aims to provide an aircraft model driving system in a wind tunnel with omnibearing, large-range and high-speed performance, wind tunnel aerodynamics and a wind tunnel virtual flight mechanical performance measuring method.
In order to solve the technical problems, the invention provides the following technical scheme:
on the one hand, provide aircraft model actuating system in wind-tunnel, including branch device, six pole parallel mechanism and linear guide, wherein:
the strut device is positioned at the rear part of the aircraft model and is used for supporting the aircraft model through a multi-component balance inside the model;
the strut device comprises a hollow tail strut main body and a rear tail strut which extends out of one end of the tail strut main body and can rotate, and a rolling driving device for driving the rear tail strut to rotate and a rolling measuring device for measuring the rotation angle of the rear tail strut are installed in the tail strut main body;
the six-rod parallel mechanism comprises six connecting rods, two ends of each connecting rod are provided with universal hinges, and the universal hinges at the front ends of the six-rod parallel mechanism are connected with the tail rod main body;
the linear guide rail is installed on the inner wall of the wind tunnel experiment section, and the universal hinge at the rear end of the six-rod parallel mechanism is connected with the slide block of the linear guide rail.
According to the invention, a rolling locking device for locking the rear tail support rod is preferably further installed in the tail support rod main body.
According to the invention, three connecting rods of the six-rod parallel mechanism are connected to the front part of the tail rod main body, and the other three connecting rods of the six-rod parallel mechanism are connected to the rear part of the tail rod main body.
According to the invention, the positions of the three connecting rods connected to the rear part of the tail rod main body on the tail rod main body are preferably positioned right behind the three connecting rods connected to the front part of the tail rod main body.
According to the invention, the front end connection points of the three connecting rods connected to the front part of the tail rod main body are positioned in the circumferential surface of the tail rod main body and are uniformly distributed, and the front end connection points of the three connecting rods connected to the rear part of the tail rod main body are also positioned in the circumferential surface of the tail rod main body and are uniformly distributed.
According to the invention, the number of the linear guide rails is 6, and 3 linear guide rails are respectively distributed on the inner walls of two sides of the wind tunnel experiment section.
According to the optimization of the invention, the aircraft model is provided with a control surface, a rotating shaft of the control surface extends to a front support rod positioned in front of the multi-component balance in the model, a control motor is arranged in the front support rod, the control motor is driven by a reduction gear to be connected with the rotating shaft of the control surface, and the control motor is controlled by a control circuit outside the wind tunnel.
On the other hand, a method for measuring the performance of an aircraft model in a wind tunnel by using the driving system is provided, and in an open-loop wind tunnel virtual flight test, the method comprises the following steps:
(a) supporting the aircraft model on a tail strut;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a six-component balance inside the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) calculating the motion position and speed of the next micro-period of the model by the dynamic equation of the aircraft according to the data similar to the real aircraft power used in the dynamic equation of the aircraft;
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) this is repeated, and if the six-bar mechanism is followed fast enough and the given time interval is small, the results of the flight dynamics can be obtained by this discretized local quasi-linear method.
In another aspect, a method for measuring performance of an aircraft model in a wind tunnel by using the above driving system is provided, where in a closed-loop wind tunnel virtual flight test with a stationary control surface, the method includes:
(a) supporting the aircraft model on a strut arrangement;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a multi-component balance inside the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) calculating the motion position and speed of the next micro-step period of the model by using data similar to the real aircraft power in a flight dynamics equation with a flight control law of the aircraft;
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) and repeating the steps, if the follow-up speed of the six-rod parallel mechanism is fast enough, and the given time interval is small, the result of the flight dynamics with the flight control law can be obtained by the discretization local quasi-linear method.
In another aspect, a method for measuring performance of an aircraft model in a wind tunnel by using the above driving system is provided, where in a closed-loop wind tunnel virtual flight test with controlled control surface, the method includes:
(a) supporting the aircraft model on a strut arrangement;
(b) giving initial deflection parameters of the control surface from predetermined flight requirements;
(c) calculating a virtual flight track with a flight control law according to the measured aerodynamic force of the aircraft model and the dynamic parameters of the aircraft given by design;
(d) comparing the virtually calculated flight trajectory with the required flight trajectory, and sending an instruction to enable the control surface to deflect according to the requirement of the control law by the obtained deviation;
(e) for example, in a flight training system, a driver can be reflected and connected into a closed loop, and then a control law result controlled by a control law is obtained according to the discretization method.
The invention has the following beneficial effects:
(a) the driving system can simultaneously provide omnibearing, large-range and high-speed motion in six degrees of freedom in a wind tunnel, has enough rigidity, strength and low interference, and can be used for measuring the high-speed aerodynamic performance of an aircraft with high maneuverability and agility in a high-speed wind tunnel;
(b) according to the driving system, through the combined change of the six-rod parallel mechanism and the rolling driving and measuring device arranged in the tail rod main body, the system can simultaneously provide large-angle motions of model pitching, yawing and rolling and the combination thereof as well as motions of three linear displacements of front and back, sinking and floating and side sliding and the combination thereof (such as conical motion and three-period motion with rolling, nutation and precession simultaneously), and measure corresponding aerodynamic performance;
(c) the invention provides a forced wind tunnel virtual flight test method for measuring the flight dynamics performance of a model, namely a forced wind tunnel virtual flight test technology, which can directly give flight dynamics data of an uncontrolled aircraft and verify the accuracy, efficiency and reliability of hardware and control program software of a control system of the controlled aircraft under flight conditions.
Drawings
FIG. 1 is a schematic overall structure diagram of an aircraft model drive system in a wind tunnel according to the present invention;
FIG. 2 is a schematic illustration of a roll driving and measuring device and aircraft model control surface configuration within an aircraft model and strut arrangement according to the present invention;
FIG. 3a is a block diagram of an open-loop wind tunnel virtual flight test;
FIG. 3b is a block diagram of the open loop control system without control laws;
FIG. 4a is a block diagram of closed loop aircraft model calculation with control laws;
FIG. 4b is a schematic diagram of a closed loop band control law;
FIG. 5a is a schematic block diagram of the closed loop control laws of a control engineer;
FIG. 5b is a functional block diagram of the flight dynamics scientist's belt control law;
FIG. 6a is a schematic diagram of driving a model in three cycles;
FIG. 6b is a schematic view of the mold being driven in a conical motion;
FIG. 7 is a schematic diagram of a transformation relationship of a coordinate axis system;
FIG. 8 is a schematic diagram of the software, hardware and driver reaction development of a controlled aircraft control system in a closed loop;
FIG. 9 is a graph of the travel time required to control the change in angle of attack for each stick in the present invention;
FIG. 10 is a position diagram of the front and rear universal hinges of the six-bar parallel mechanism of the present invention.
Detailed Description
In order to make the technical problems, technical solutions and advantages of the present invention more apparent, the following detailed description is given with reference to the accompanying drawings and specific embodiments.
The invention provides an aircraft model driving system in a wind tunnel, which comprises a strut device 103, a six-rod parallel mechanism 106 and a linear guide rail 112, as shown in figures 1-2, wherein:
the strut device 103 is positioned at the rear part of the aircraft model 101 and is used for supporting the aircraft model 101 through a multi-component balance 102 inside the model;
the strut device 103 comprises a hollow tail bar main body 1031 and a rear tail strut 1032 which extends from one end of the tail bar main body 1031 and is rotatable (the rear tail strut 1032 can be supported in the tail bar main body 1031 through a front bearing and a rear bearing 206 in particular), and the front end of the rear tail strut 1032 supports the aircraft model 101 through a multi-component balance 102 in the interior of the model;
the tail rod main body 1031 is internally provided with a rolling driving device 207 for driving the rear tail strut 1032 to rotate and a rolling measuring device 208 for measuring the rotation angle of the rear tail strut 1032; a rolling locking device 209 for locking the rear tail strut 1032 can be further installed in the tail rod main body 1031 so as to improve the firmness of the rear tail strut 1032;
the six-rod parallel mechanism 106 comprises six connecting rods, two ends of each connecting rod are respectively provided with a universal hinge (specifically, a spherical hinge), the universal hinge 107 at the front end of the six-rod parallel mechanism 106 is connected with the tail rod main body 1031, and the radial position and the azimuth angle of the specific connection can be adjusted according to requirements;
the linear guide rail 112 is installed on the inner wall of the wind tunnel experiment section 109 (the specific up-down and front-back positions can be changed according to the experiment requirement), and the universal hinge 108 at the rear end of the six-rod parallel mechanism 106 is connected with the slide block of the linear guide rail 112.
The position of the linear guide rail 112 on the wind tunnel side wall and the position of the front universal hinge on the tail rod main body 1031 can be adjusted according to the test requirements and specificity, so as to meet the optimal test requirements of comprehensive moving speed, rigidity and the like.
The driving system can simultaneously provide omnibearing, large-range and high-speed motion in six degrees of freedom in a wind tunnel, has enough rigidity, strength and low interference, and can be used for measuring the high-speed aerodynamic performance of an aircraft with high maneuverability and agility in a high-speed wind tunnel;
if the driving system enables the rolling angular speed and the nutation angular speed to be synchronous, the coning motion mode of the aircraft model can be obtained;
the driving system can simultaneously provide large-angle motion of model pitching, yawing and rolling and combination thereof and motion of three linear displacements of front and back, sinking and floating and side slipping and combination thereof (such as conical motion and three-period motion with rolling, nutation and precession at the same time) through the combined change of the six-rod parallel mechanism and a rolling driving and measuring device arranged in the tail rod main body, and measures corresponding aerodynamic performance;
the driving system can not only measure the aerodynamic force of the model, but also be used for the virtual flight of the aircraft model, and realizes the integrated research of the aerodynamic force and the flight mechanics to evaluate the flight control quality.
In the invention, the two ends of the connecting rods of the six-rod parallel mechanism positioned in the wind tunnel are connected by the universal hinges, and all the connecting rods are in a two-force rod state, so that the stress form can bear larger aerodynamic force and higher rigidity than the traditional cantilever beam, thereby enabling the test to be carried out at high speed even at supersonic speed; the model also has higher position and attitude control precision; under the action of the same aerodynamic force, each support rod can be thinner than the section of the cantilever beam so as to further reduce the blockage degree of the wind tunnel; with the struts and aft tail struts all behind the aircraft model, strut interference will be less than that supported by the side struts.
In the present invention, three links of the six-bar parallel mechanism 106 are connected to the front portion of the tail bar main body 1031, and the other three links of the six-bar parallel mechanism 106 are connected to the rear portion of the tail bar main body 1031, in order to support the aircraft model 101 more smoothly. Further, to reduce disturbance of the airflow, the three links connected to the rear portion of the tail bar main body 1031 are preferably located directly behind the three links connected to the front portion of the tail bar main body 1031 on the tail bar main body 1031. In addition, for convenience of control and position calculation, the front end connection points of the three links connected to the front portion of the tail bar main body 1031 may be located in the circumferential surface of the tail bar main body 1031 and uniformly distributed, and the front end connection points of the three links connected to the rear portion of the tail bar main body 1031 are also located in the circumferential surface of the tail bar main body 1031 and uniformly distributed.
The number of the linear guide rails 112 may be 6, and 3 linear guide rails are respectively distributed on the inner walls of two sides of the wind tunnel experiment section 109, so that one linear guide rail drives and controls one connecting rod.
In conclusion, the invention can be used for carrying out three-period motion and cone motion tests besides large-angle dynamic tests. At the moment, the six-rod parallel mechanism drives the fuselage axis of the aircraft model to do conical rotation around the wind tunnel axis, and meanwhile, the motor in the support rod device drives the model to do rotary motion. This results in three cycles of movement of the model aircraft (fig. 6a, where α is the angle of attack and β is the angle of sideslip). When the model is driven by the motor in the strut device at the same rotational angular velocity as in the three-cycle motion, the six-strut parallel mechanism keeps the nutation angle unchanged, and only the precession angle changes, which results in the coning motion of the model around the wind tunnel axis (fig. 6b, wherein
Is the cone motion angular velocity). The cone angle of the model can be easily adjusted by changing the nutation angle in a wind tunnel blowing test through a six-rod parallel mechanism.
The invention can carry out omnibearing, large-range, high-speed, small-interference, high-rigidity and high-precision wind tunnel aerodynamic tests and carry out 'forced' wind tunnel flight dynamic tests by the device (namely a driving system) as two basic points of the invention. The first basic point, namely the motion space, the kinematic and dynamic properties of the device, will be described.
The six-rod parallel mechanism adopted by the scheme is shown in figure 1: according to the characteristics and requirements of the wind tunnel test, six linear guide rails are arranged on different circumferential positions of the side wall of the wind tunnel test section in parallel with the longitudinal axis of the wind tunnel. The upper and lower positions of the guide rail can be adjusted according to the use requirement. The sliding block moving pair (P) on the linear guide rail is connected with the connecting rod through a rear universal hinge (U) or a spherical hinge (S). The front ends of the connecting rods are connected with a supporting rod device for supporting the model by a front universal hinge (U) (figure 2).
The topological structure of the mechanism is characterized by 6-SOC { -P-S-U } PM
Two front and back lines of universal hinge pairs, three universal hinge pairs in each line, each other becomes 120 degrees:
thus, the whole mechanism consists of six parallel sliding pairs, namely a universal hinge pair, a parallel link rod and a universal hinge pair, namely a 6-PUU mechanism.
The topological structure of the mechanism is characterized in that:
number of pairs of motion: m is 18; a spherical pair (ms)6, a universal hinge pair (mu)6 and a moving pair (mp)6
Number of members (n): 14
Number of branches (n)b)：6
Simple number of branches (n)bs)：6
Moving platform poc (mpa): mpa ═ t3 r3]T
POC set dimension: 6
Number of independent loops: v is 5
Number of independent displacement equations: 30
Degree of freedom: 6
Local degree of freedom: 6
Number of overconstrained: 0
Redundancy: 0
A driving pair: 6 p pairs
BKC number (nBKC): 1
BKC coupling degree: 3
The type of degree of freedom: full DOF
Motion input and output decoupling number: there is no decoupling.
It should be pointed out that the parallel mechanism is a novel parallel mechanism, and is characterized in that:
in order to reduce the blockage degree of the wind tunnel test, the strut device and the linear guide rail are arranged in the downstream direction;
the six universal hinges are two rows and are fixed on the tail rod main body at the rear part of the model in a downstream mode. Thus, a so-called frontal plane of motion may be defined as a virtual plane;
the universal hinge at the rear part of the connecting rod is connected on the sliding block of the linear guide rail. The positions of the rear universal hinges on the linear guide rails also change along with the change of the orientation of the model, and are not on the same plane, and can be actually defined as another virtual plane.
The parallel mechanism has six degrees of freedom and six control quantities. Because the parallel mechanism is easy to solve reversely (the position of the slide rail on the linear motion guide rail is solved by inputting the motion parameters of the model), the position of the slide rail on the linear motion guide rail is solved by solving the equation reversely after the geometric position and the orientation of the model are given, and the motion space, the kinematics and the dynamic performance, the sensitivity, the singularity, the error analysis and the like of the mechanism are further analyzed.
The model has three linear variables, three angular variables, which are six independent variables. The positions of the sliding blocks on the six linear motion guide rails are dependent variables. And the positions of the upper and lower positions of the guide rail and the universal hinge on the tail rod main body are parameters. Therefore, in the following equation derivation, six dependent variables (six slider positions on the guide rail) are found given the parameters according to the given position of the model.
To this end, two coordinate systems can be established: a moving coordinate system moving with the model and a fixed coordinate system fixed on the wind tunnel test section (see fig. 7, wherein θ, ψ, and γ are the pitch angle, yaw angle, and roll angle of the model, respectively). The idea of solving the problem is as follows:
firstly, the position and the azimuth angle (intermediate variable) of the front universal hinge on the front dynamic plane are obtained according to the position and the azimuth angle (independent variable, which can be any function of a given value or time) of the model and the position (parameter variable) of the front universal hinge on the tail rod main body.
And (4) solving the position and the azimuth angle of the intermediate variable on the front dynamic plane on the fixed coordinate system of the rear wind tunnel test section through matrix conversion.
Then, the position (dependent variable) of the slide block on the moving linear motion guide rail is obtained under the condition that the vertical, horizontal and connecting rod lengths of the motion guide rail in the wind tunnel are counted (all parameters).
The second essential point of the invention, namely the essential point of the "forced" wind tunnel flight dynamics test (wind tunnel virtual flight test) with this drive system, is explained below.
Firstly, in order to ensure the success of the closed-loop wind tunnel virtual flight test, the rapid following performance of the system is the key of the success of the test. From the analysis of the time chain, the time chain may be composed of the following parts: time required for aerodynamic force measurement (Δ t)1) Time required for flight dynamics equations and control law calculations (Δ t)2) Time required for the parallel mechanism to go into position (Δ t)3) …, in view of the fast response time (Δ t) of the mechanism3) Is the longest time period required, and therefore, the time (delta t) required for the parallel mechanism to be in place is shortened3) Is one of the key points of the invention. To verify and illustrate this, and for the practical dimensions of the case proposed by the invention, fig. 9 shows the speeds and ranges of movement required for the parallel links of the parallel mechanism in order to simulate the controlled movements of the aircraft in the case of a "forced" wind tunnel virtual flight test. If the maximum rod speed of the linear guide is chosen to be 1 m/s, the calculations show that the time required for the rods 1 to 6 to reach angles of attack from-30 degrees to +30 degrees is approximately 0.05 seconds, 0.12 seconds, 0.06 seconds, 0.14 seconds, 0.19 seconds and 0.09 seconds, respectively. The angular velocity ranges from 300 degrees/second to 1200 degrees/second. The angular velocity can be increased further if the rod speed is increased and the rod position is optimized. Therefore, the driving system provided by the invention can sufficiently meet the requirement of a forced wind tunnel virtual flight test on the running speed of the mechanism.
In order to realize the forced wind tunnel virtual flight test, the aircraft model is required to be provided with a control surface and a corresponding deflection control mechanism, and the structure is as follows:
as shown in fig. 2, the aircraft model 101 has a control surface 201 (the embodiment shown in the figure is embodied as an elevator), a rotating shaft 202 of the control surface 201 extends to a front support rod 203 at the front part of the model internal multi-component balance 102, a control motor 205 is arranged in the front support rod 203, the control motor 205 drives the rotating shaft 202 connected with the control surface 201 through a reduction gear 204, and the control motor 205 is controlled by a control circuit outside the wind tunnel.
The forced wind tunnel virtual flight test can be further divided into an open-loop wind tunnel virtual flight test (XN-1 type) and a closed-loop wind tunnel virtual flight test. The latter can be further divided into two cases of adding the control law (XN-2 type) to the flight dynamics equation only by the model control surface fixation and adding the control law of the model control surface or the human-machine control (XN-3). That is, the wind tunnel virtual flight test can be divided into three different forms and stages:
(a) selecting a plurality of fixed control plane deflection angles, carrying out a flight trajectory comparison study, and qualitatively evaluating the influence of flight dynamics (XN-1, figures 3a and 3 b); wherein each process in fig. 3b is as follows:
(b) under the condition that the control surface of the wind tunnel model is fixed, a flight control law module is added into a flight dynamics motion equation of the model, the measured parameters and the aircraft control module are combined to form a virtual flight system, and the influence of the control law is researched (XN-2 type, figures 4a and 4 b);
(c) a control surface control system is added to the model to realize real-time control (controlled by the control surface control system or a pilot) so as to more directly verify the accuracy, efficiency and reliability of the hardware, the pilot and the control program software of the control system of the controlled aircraft under flight conditions (type XN-3, fig. 5a, 5 b).
In an open-loop wind tunnel virtual flight test (XN-1, FIGS. 3a and 3b), the method for measuring the performance of an aircraft model in a wind tunnel may include the steps of:
(a) supporting the aircraft model on a strut arrangement;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a multi-component balance (specifically, a six-component balance) in the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) data for real aircraft dynamics similarities (mass, center of mass, moments of inertia, etc. are similar) used in the dynamics equations for the aircraft (without the need for a model to model the real aircraft dynamics similarities). Calculating the position and velocity of motion for the next microstep period (e.g., 0.001 second) of the model from the aircraft's equations of dynamics;
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) this is repeated again and again, if the six-bar parallel mechanism is followed fast enough and the given time interval (i.e. microstep period) is small (e.g. 0.001 seconds), the results of the flight dynamics can be obtained in this discretized local quasi-linear method.
The second type of the forced wind tunnel virtual flight test is a closed-loop wind tunnel virtual flight test, and can be divided into two methods of control surface immobilization and control surface control (fig. 4a and 4b and fig. 5a and 5 b).
For a closed-loop wind tunnel virtual flight test (XN-2, fig. 4a and 4b) with a fixed control surface, the control surface of an aircraft model is fixed, a control law is only added into a flight dynamics equation, and a result with the control law is obtained according to the discretization method. In fact, the main difference between the closed-loop wind tunnel virtual flight test with the unmoved control surface and the open-loop wind tunnel virtual flight test with the unmoved control surface is that the former adds a flight control law in the flight dynamics calculation, and the aircraft model reaches a new model orientation according to the result. Specifically, in a closed-loop wind tunnel virtual flight test with a stationary control surface, the method for measuring the performance of an aircraft model in a wind tunnel may include the steps of:
(a) supporting the aircraft model on a strut arrangement;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a multi-component balance (specifically, a six-component balance) in the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) the data of real aircraft dynamics similarity (mass, centroid, moment of inertia, and the like are similar) used in the flight dynamics equation with flight control law of the aircraft calculates the motion position and speed of the model for the next micro-step period (for example, 0.001 second);
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) this is repeated, and if the six-bar parallel mechanism is followed fast enough and the given time interval (i.e. microstep period) is small (e.g. 0.001 seconds), the result of the flight dynamics with the flight control laws can be obtained by the discretized local quasi-linear method.
In a closed-loop wind tunnel virtual flight test (XN-3, figures 5a, 5b and 8) with a control surface controlled by a control law, the control surface changes an angle due to the change of the control law during the wind tunnel test, and control software, hardware or a driver in a flight training system is switched into the closed loop. The method for measuring the flight dynamics performance of the aircraft model in the wind tunnel can comprise the following steps:
(a) supporting the aircraft model on a strut arrangement;
(b) giving initial deflection parameters of the control surface from predetermined flight requirements;
(c) calculating a virtual flight path (position, orientation and motion parameters of the aircraft) with a flight control law according to the measured aerodynamic force of the aircraft model and the dynamic parameters (such as mass, mass center and moment of inertia) of the aircraft given by design;
(d) comparing the virtually calculated flight trajectory with the required flight trajectory, and sending out a command to enable a control surface (the mechanism of which is shown in figure 2) to deflect according to the requirement of a control law (such as a PID (proportion integration differentiation) module) on the obtained deviation;
(e) the driver reflection can also be switched into a closed loop, as in a flight training system. And then obtaining a control law result of the control surface controlled by the control law according to the discretization method.
Therefore, by the closed-loop wind tunnel virtual flight test, not only can instantaneous unsteady and nonlinear aerodynamic data be obtained, but also the accuracy, efficiency and reliability of the hardware and control program software of the control system of the controlled aircraft under the flight condition can be checked; or to evaluate the operation of a pilot in training.
In the forced wind tunnel virtual flight test of the closed loop control law, the key point is to spread a transfer function module of a control system of a control surface to form an algorithm and realize the real-time control of a computer. The invention can be realized by the following technical scheme, for example, for a linear steady system:
x(t)＝AX+BU
y(t)＝CX+DU
if the reference motion is horizontal linear flight, the dimensionless matrix expression of the longitudinal small disturbance equation set can be written as follows:
expression of equation of state
In the formula:
a and B are matrix coefficients of a longitudinal dimensionless equation of motion.
The discretization equation is as follows:
x(k+1)＝GX(k)+HU(k)
y(k+1)＝CX(k)+DU(k)(k＝0,1,2…)
wherein G, H, C, D are constant matrixes of longitudinal dimensionless equation of motion, and
G＝eAT
therefore, the state space equation of the continuous system can be discretized, and the test result with the closed loop control law can be obtained through computer iteration and corresponding tracking of the parallel mechanism.
If the given time interval is small, for example 0.001 seconds, extrapolation of the measured values from the previous instants can be used to determine new control surface position and motion parameter values.
While the foregoing is directed to the preferred embodiment of the present invention, it will be understood by those skilled in the art that various changes and modifications may be made without departing from the spirit and scope of the invention as defined in the appended claims.
Claims (5)
1. The method for measuring the performance of the aircraft model in the wind tunnel by using the aircraft model driving system in the wind tunnel is characterized in that the aircraft model driving system in the wind tunnel comprises a support rod device, a six-rod parallel mechanism and a linear guide rail, wherein:
the strut device is positioned at the rear part of the aircraft model and is used for supporting the aircraft model through a multi-component balance inside the model;
the strut device comprises a hollow tail strut main body and a rear tail strut which extends out of one end of the tail strut main body and can rotate, and a rolling driving device for driving the rear tail strut to rotate and a rolling measuring device for measuring the rotation angle of the rear tail strut are installed in the tail strut main body;
the six-rod parallel mechanism comprises six connecting rods, two ends of each connecting rod are provided with universal hinges, and the universal hinges at the front ends of the six-rod parallel mechanism are connected with the tail rod main body;
the linear guide rail is arranged on the inner wall of the wind tunnel experiment section, and a universal hinge at the rear end of the six-rod parallel mechanism is connected with a sliding block of the linear guide rail;
the number of the linear guide rails is 6, and 3 linear guide rails are distributed on the inner walls of the two sides of the wind tunnel experiment section respectively;
the aircraft model is provided with a control surface, a rotating shaft of the control surface extends to a front support rod positioned in front of the multi-component balance in the model, a control motor is arranged in the front support rod, the control motor is driven by a reduction gear to be connected with the rotating shaft of the control surface, and the control motor is controlled by a control circuit outside the wind tunnel;
in an open-loop wind tunnel virtual flight test, the method comprises the following steps:
(a) supporting the aircraft model on a tail strut;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a six-component balance inside the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) calculating the motion position and speed of the next micro-period of the model by the dynamic equation of the aircraft according to the data similar to the real aircraft power used in the dynamic equation of the aircraft;
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) the operation is repeated, if the follow-up speed of the six-rod mechanism is fast enough, and the given time interval is very small, the result of the flight dynamics can be obtained by the discretization local quasi-linear method;
in a closed-loop wind tunnel virtual flight test with a motionless control surface, the method comprises the following steps:
(a) supporting the aircraft model on a strut arrangement;
(b) deflecting a control surface of the aircraft model to a fixed deflection angle;
(c) when a control surface of the aircraft model deflects, a multi-component balance inside the model measures instantaneous aerodynamic force and aerodynamic moment;
(d) calculating the motion position and speed of the next micro-step period of the model by using data similar to the real aircraft power in a flight dynamics equation with a flight control law of the aircraft;
(e) inputting the result into a six-rod parallel mechanism to enable the aircraft model to move to a new position at the calculated new speed;
(f) the operation is repeated, if the follow-up speed of the six-rod parallel mechanism is fast enough, and the given time interval is small, the result of the flight dynamics with the flight control law can be obtained by the discretization local quasi-linear method;
in a closed-loop wind tunnel virtual flight test with a controlled control surface, the method comprises the following steps:
(a) supporting the aircraft model on a strut arrangement;
(b) giving initial deflection parameters of the control surface from predetermined flight requirements;
(c) calculating a virtual flight track with a flight control law according to the measured aerodynamic force of the aircraft model and the dynamic parameters of the aircraft given by design;
(d) comparing the virtually calculated flight trajectory with the required flight trajectory, and sending an instruction to enable the control surface to deflect according to the requirement of the control law by the obtained deviation;
(e) for example, in a flight training system, a driver can be reflected and connected into a closed loop, and then a control law result controlled by a control law is obtained according to the discretization method.
2. The method of claim 1, wherein a roll lock is also mounted in the tail boom body for locking the rear tail boom.
3. The method of claim 1, wherein three links of the six-bar parallel mechanism are connected at a front portion of the tail bar body and three other links of the six-bar parallel mechanism are connected at a rear portion of the tail bar body.
4. The method of claim 3, wherein the three links connected to the rear of the tail bar body are positioned directly behind the three links connected to the front of the tail bar body on the tail bar body.
5. The method of claim 3, wherein the front end connection points of the three links connected to the front portion of the tail boom body are located within and evenly distributed about the circumference of the tail boom body, and the front end connection points of the three links connected to the rear portion of the tail boom body are also located within and evenly distributed about the circumference of the tail boom body.
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| CN202010802116.2A CN111929023B (en) | 2020-08-11 | 2020-08-11 | Aircraft model driving system in wind tunnel and performance measuring method |
| PCT/CN2021/123032 WO2022033608A1 (en) | 2020-08-11 | 2021-10-11 | Drive system and performance measurement method for aircraft model in wind tunnel |
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| CN111929023B (en) * | 2020-08-11 | 2021-08-06 | 日照坤仑智能科技有限公司 | Aircraft model driving system in wind tunnel and performance measuring method |
| CN112525474B (en) * | 2020-12-22 | 2021-08-17 | 中国空气动力研究与发展中心高速空气动力研究所 | Method for obtaining blocking interference factor of transonic wind tunnel |
| CN114112281A (en) * | 2021-07-28 | 2022-03-01 | 上海华模科技有限公司 | Wind tunnel test device for decoupling measurement of aerodynamic dynamic derivative |
| CN114441133B (en) * | 2021-12-28 | 2023-06-06 | 中国航天空气动力技术研究院 | Pose calibration method and equipment for attack angle-double-rotating-shaft mechanism for wind tunnel test |
| CN114486167B (en) * | 2022-04-18 | 2022-06-24 | 中国空气动力研究与发展中心低速空气动力研究所 | Follow-up sealing system and method for floor notch of low-speed wind tunnel model tail support ground effect test |
| CN114563153B (en) * | 2022-04-28 | 2022-07-01 | 中国空气动力研究与发展中心超高速空气动力研究所 | Ultrahigh-speed pneumatic test device for accelerating gas test through track |
| CN114608794B (en) * | 2022-05-11 | 2022-07-19 | 中国航空工业集团公司哈尔滨空气动力研究所 | Method for measuring aerodynamic coefficient of model wind tunnel virtual flight test |
| CN114993606B (en) * | 2022-05-31 | 2023-03-21 | 中国科学院力学研究所 | Wind tunnel test result processing method for unsteady pressure and aerodynamic data |
| CN115186376B (en) * | 2022-06-21 | 2022-12-20 | 中国航空工业集团公司哈尔滨空气动力研究所 | Small-aspect-ratio flying wing aircraft rolling torque model and parameter identification method thereof |
| CN115343012B (en) * | 2022-07-07 | 2023-04-07 | 中国航空工业集团公司哈尔滨空气动力研究所 | Unsteady-state large-amplitude oscillation test method |
| CN114923657B (en) * | 2022-07-21 | 2022-10-04 | 中国空气动力研究与发展中心高速空气动力研究所 | Vertical rod rope supporting system for full-aircraft flutter wind tunnel test |
| CN115077846B (en) * | 2022-07-28 | 2022-10-28 | 中国空气动力研究与发展中心高速空气动力研究所 | Large wind tunnel ground preparation platform |
| CN115717981B (en) * | 2023-01-09 | 2023-04-21 | 中国空气动力研究与发展中心超高速空气动力研究所 | Resident chamber device for multiple spray pipes in wind tunnel test and use method |
| CN116062183B (en) * | 2023-03-22 | 2023-06-09 | 中国空气动力研究与发展中心设备设计与测试技术研究所 | High-speed towing system for cross-medium aircraft water-air crossing test |
| CN116399543B (en) * | 2023-04-10 | 2023-12-12 | 四川省机械研究设计院(集团)有限公司 | Six-degree-of-freedom wind tunnel model supporting system based on Hexaglide parallel mechanism and control method |
| CN116380397B (en) * | 2023-05-31 | 2023-08-18 | 成都流体动力创新中心 | Typical maneuvering course simulation test device based on magnetic levitation flight wind tunnel |
| CN116878819B (en) * | 2023-09-06 | 2024-02-06 | 中国人民解放军32806部队 | Wind tunnel test device and method for aerodynamic characteristics of aircraft |
| CN116933400B (en) * | 2023-09-13 | 2023-11-21 | 中国航空工业集团公司哈尔滨空气动力研究所 | Method for constructing aerodynamic moment model of coupling uncertainty |
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| DE102009002392A1 (en) * | 2009-04-15 | 2010-11-04 | Airbus Deutschland Gmbh | System and method for determining local accelerations, dynamic load distributions and aerodynamic data in an aircraft |
| CN104931222B (en) * | 2015-05-04 | 2018-04-10 | 中国航天空气动力技术研究院 | Machine bullet separated track captive test system |
| CN105784318B (en) * | 2016-03-04 | 2017-07-18 | 中国空气动力研究与发展中心低速空气动力研究所 | A kind of low-speed wind tunnel model flight experimental system and method |
| CN106289708B (en) * | 2016-07-26 | 2018-08-07 | 中国航天空气动力技术研究院 | The pose scaling method of motion for captive trajectory wind tunnel test |
| CN109668710B (en) * | 2018-12-18 | 2020-04-07 | 大连理工大学 | Multi-dimensional vibration control method for strut tail support type aircraft model |
| CN110160730B (en) * | 2019-06-24 | 2021-10-29 | 日照坤仑智能科技有限公司 | Device and method for testing separation performance of aircraft plug-in high-speed wind tunnel |
| CN111473949B (en) * | 2020-04-14 | 2022-04-01 | 日照坤仑智能科技有限公司 | Device and method for measuring time difference of falling wings, rolling and washing of aircraft |
| CN111929023B (en) * | 2020-08-11 | 2021-08-06 | 日照坤仑智能科技有限公司 | Aircraft model driving system in wind tunnel and performance measuring method |
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