CN114912202A - Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body - Google Patents
Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body Download PDFInfo
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Abstract
The invention belongs to the technical field of airplane control, and relates to a body propulsion integrated coupling control method for a wide-speed-range air-breathing power aircraft. The invention establishes an integrated dynamic model from the aspect of aeromechanics/engine thermodynamic cycle coupling, lists state space equations by using disturbance linearization of control differential equations, and provides an integrated coupling feedback control strategy of a wide-speed-domain aircraft under a large-range Mach number by adopting a pole allocation method. The invention obviously improves the control quality and the operation stability of the aircraft in the aspect of coping with the flying/pushing coupling effect, has the characteristics of smaller overshoot, higher response speed and fast convergence, and has wide application prospect.
Description
Technical Field
The invention belongs to the technical field of aircraft control, and relates to a body propulsion integrated coupling control method for a wide-speed-range air-breathing power aircraft, which is suitable for wide-speed-range air-breathing power aircraft with remarkable body propulsion integrated characteristics.
Background
The wide-speed range air-breathing power aircraft has the maximum Mach number of more than 5, uses a sub-combustion/super-combustion ramjet engine and combined power thereof, has the characteristic of repeated use in horizontal take-off and landing, can fly in a maneuvering way in a near space with the altitude of 20-100 kilometers, and depends on rare atmosphere to execute tasks such as striking, investigation, transportation, confrontation and the like. The aircraft has the characteristics of high flying speed and response speed, large flying envelope, large lift-drag ratio, large specific impulse and the like, has remarkable advantages compared with the conventional aircraft, has wide development prospect, and has great value in the military and civil fields.
The air-breathing wide-speed-range aircraft has the characteristics of multiple modes, multiple variables, nonlinearity and time-varying property, and brings difficulty to a control system of the air-breathing wide-speed-range aircraft. The new generation of wide speed range aircraft typically requires compression of the air entering the ram engine by the fuselage forebody and air intake, so aerodynamic forces over a wide range of mach numbers, coupling and interplay of propulsion system operating conditions are mainly manifested in the change of air mass flow through the air intake compression angle at flight attitude, thereby affecting thrust. Meanwhile, the thrust influences the posture in two aspects. Firstly, due to the special structure of the combined power engine, the thrust center cannot be designed at the mass center of the airplane, so that the eccentric moment is generated. Second, the high velocity expanding gas flow from the jets has an asymmetric pressure on the asymmetric outer jets, creating additional torque. Therefore, coordinated control of attitude, speed, and thrust is particularly important. Due to the remarkable characteristic of the body propulsion integration of the air-breathing wide-speed-range aircraft, the traditional flight control/engine control decoupling design strategy is not applicable any more.
Extensive research has been conducted by many scholars on the control of wide speed range aircraft over a wide mach number range. HEJ, QI R and the like predict the sliding mode surface combining the selected tracking error and the high-order derivative by using a feedback linear control decoupling method, and the sliding mode surface is incorporated into a quadratic performance index of prediction control on the basis. And the effect of reducing buffeting caused by switching of the sliding mode control law is achieved by using the minimized performance index, so that the problem of online calculation of predictive control is solved, and the predictive sliding mode control law with an explicit solution is finally obtained. Students such as the Li navy and Huang Yinling and the like compensate the uncertainty of the reentry maneuver nonlinear model of the wide-speed-range aircraft on line based on the generalized fuzzy neural network, and improve the linear control performance of the trajectory. The track linear control method for Liukai, Guokai and the like solves the problem of track tracking guidance of the climbing section of the air-breathing combined power wide-speed-range aircraft, and provides a guidance parameter design criterion which can give consideration to both closed-loop stability and control quality. X.w. and q.qi have made a series of original contributions that propose a fuzzy optimal control strategy to address the problem of tracking control of hypersonic aircraft affected by unknown dynamics. Recently, they have been working on applying a simplified limited time fuzzy neural controller with a predetermined performance to a passenger aircraft, which guarantees excellent real-time performance with a simple structure. Bu provides a funnel type non-affine controller based on neural network approximation, can help the wide speed domain aircraft of air-breathing type to realize the ability of tracking predetermined orbit. For both methods, the required transient and steady state performance can be guaranteed. The problem of fly/push coupling control of air-breathing wide speed range aircraft has not been solved by adding a feedback loop.
Disclosure of Invention
The air-breathing wide-speed-range aircraft has obvious flying/pushing coupling effect in a large flying envelope, and the control effect of the traditional decoupling control is affected by the coupling effect and is poor. The invention designs an aircraft/engine coupling control method, which avoids complex controller parameter adjustment and flight/propulsion system decoupling, and skillfully introduces a new feedback channel on the basis of a traditional negative feedback model by designing state variables mainly through the basic theory of pole allocation to form a coupling feedback control system.
The technical scheme of the invention is as follows:
the integrated coupling control method for the propulsion of the wide-speed-range air-breathing power aircraft body specifically comprises the following steps:
(1) aircraft dynamics modeling
(1.1) pneumatic binning modeling
Dividing the surface of the aircraft into a plurality of small areas by using a surface element method, replacing the small areas with parallelograms or triangles, and calculating aerodynamic force and moment on a surface element;
(1.2) supersonic pneumatic computation
Calculating the pressure and pressure coefficient of each surface element by adopting an ultrasonic pneumatic estimation method, and adding the vectors of each surface element to obtain the total aerodynamic force and the aerodynamic moment under the working condition;
(1.3) Propulsion System calculations
In the thrust estimation, the scramjet engine is divided into an air inlet channel, an isolator, a combustion chamber and a spray pipe;
in the air inlet part, describing flow field parameters by adopting shock waves, expansion waves and an intersection theory thereof, and calculating all shock waves and expansion waves generated in a plane due to the rotation of an inlet wall and the intersection of the shock waves and the expansion waves; the waves divide a two-dimensional plane area into a plurality of small areas, the internal airflow parameters in each small area are the same and are determined by shock waves or expansion waves in the front area;
calculating airflow parameters of the isolator and the combustion chamber by adopting a quasi-one-dimensional model, obtaining a basic conservation equation of the airflow in the isolator and the combustion chamber without considering airflow friction, fuel injection, length change of a pre-combustion shock wave string and reaction rate factors, and solving by combining an empirical formula;
the plume modeling is adopted for the tail nozzle, a two-dimensional model of the tail nozzle is divided into small areas along a body axis, and the plume model is built on the basis of equal air pressure above and below the shear layer;
(2) analysis of dynamics characteristics of wide-speed-range air-breathing dynamic aircraft
(2.1) establishing a wide-speed-domain flight dynamics model:
in the formula: m is the pitching moment, ω z Is the pitch angular velocity, J z Is the pitch moment of inertia, V is the velocity, theta is the track inclination, H is the flight altitude, m is the aircraft mass,
p, X, Y respectively represents thrust, drag and lift of the engine, g represents gravity acceleration, and alpha represents an attack angle;
the above is a classical aircraft dynamics model, however, due to the existence of obvious fly-push coupling characteristics, the wide-speed-range air-breathing power aircraft needs to add description of an actual fuel equivalence ratio, the fuel equivalence ratio is related to speed, height and fuel equivalence ratio commands, and the expression is as follows:
wherein the content of the first and second substances,
refers to the command of the fuel equivalence ratio,
is a fuel equivalence ratio command.
(2.2) performing linearization expansion on the aircraft in a certain flight state according to a small disturbance linearization theory; the expansion form is that each line of the equation is formed
In the form of (a);
(2.3) in order to make the expression after the flight dynamics formula linearization easy to write, a simplified symbol of an equation coefficient is introduced; the static stability and the static maneuverability of the airplane are important reference indexes for evaluating the stability of the airplane, and the specific indexes are as follows:
a 24 the static stability of the airplane is represented, and is the acceleration deflection of the rotation angle of the airplane around the body axis caused by the change of a unit attack angle; wherein the content of the first and second substances,
for changes in pitch moment due to aerodynamic forces resulting from changes in unit angle of attack,
change in pitching moment due to thrust resulting from change in unit angle of attack, J z Is the moment of inertia.
a 25 The characteristic of static maneuverability, in particular the efficiency of elevators, is the O-turn of the aircraft caused by the unit deflection of the steering mechanism z1 The sign of the shaft rotation angle acceleration deviation depends on the aerodynamic layout of the airplane and is a negative value for a normal airplane; wherein the content of the first and second substances,
is caused by unit deflection of steering mechanism z1 The moment offset of the shaft is determined by the amount of moment offset,
is caused by unit deflection of steering mechanism z1 Moment coefficient offset of axis, q is the dynamic pressure of the oncoming airflow, S is the aircraft reference area, b A Is the mean aerodynamic chord length.
(2.4) establishing a control-oriented aircraft/propulsion integrated coupling state space equation:
from the aircraft/propulsion integrated coupling state space equation, the system has two control quantities which are respectively an elevator and a fuel equivalence ratio instruction; the state quantities are four, namely fuel equivalence ratio, angular velocity, pitch angle and attack angle;
(3) coupled feedback control based on pole allocation
Coupling feedback control is realized through pole allocation of a multi-input multi-output system, the control process comprises the steps of selecting an expected pole position according to expected overshoot and adjusting time, solving controller parameters and controlling distribution, and the specific process comprises the following steps:
(3.1) desired pole selection
According to longitudinal short-period disturbing motion
The system state space equation after the control signal is added is obtained from the time domain perspective as follows:
if the value of the K matrix is such that A + BK forms a matrix that stabilizes the system asymptotically, with x (0) ≠ 0, then
The eigenvalues of A + BK are the poles of the regulator; if the pole of the regulator is located in the left half-plane of the coordinate system, then
The problem of placing the poles of the regulator in the desired positions is called pole placement;
the process of desired pole selection is as follows:
according to the logic of the elevator control attitude and the fuel equivalence ratio control system, the four-order system is divided into a three-order system and a first-order system which respectively correspond to the attack angle control system and the fuel equivalence ratio control system. The three-order system is provided with two dominant poles and a non-dominant pole, and the dynamic index of the step response of the system is mainly determined by the two dominant poles; the position of the dominant pole has a clear corresponding relation with the performance index of the corresponding system, and the dominant pole can be selected by calculating the damping ratio and the natural frequency; the formula is as follows:
wherein s is 1,2 Representing two dominant poles, xi is the system damping ratio, omega n Is the natural frequency of the system, t s To adjust the time, σ is the system overshoot.
After the selection of the two dominant poles is completed, the third pole, i.e. the non-dominant pole, can be autonomously selected according to the real part of the dominant pole, and a value more than 5 times that of the dominant pole is usually selected as the non-dominant pole, i.e.:
s 3 ＝-nξω n
n is a selected multiple that can be adjusted by the actual control effect.
The first order system has one pole, noted:
s 4 ＝-1/T
where T is the first order system time constant and the first order system settling time is typically 4T, the poles of the first order system can be selected by the ideal settling time.
To this end, the positions of the four desired poles can all be determined:
(3.2) control parameter solving
(3.3) control of dispensing
The logic of the control matrix is:
when calculating the adjustment quantity delta of the elevator z In the process of (a), k 11 Δω z And k 12 Δ α is a differential term and a proportional term, respectively;
is a coupled control item of the flight/propulsion system,
expressed fuel equivalence ratio error integral term and liftThe rudder adjustment is irrelevant, so k 14 0; similarly, the fuel equivalence ratio command adjustment is solved
In the course of (2), the process,
and
respectively representing a proportional term and an integral term, k 22 Δ α is the coupling term, k 21 Δω z Is an unrelated item;
(4) controller dynamic adjustment
In the wide-speed-range air-breathing power aircraft, pneumatic parameters may significantly change in the wide-speed-range flight process, and the control matrix K is adjusted through cyclic detection, so that the element values of the A + BK matrix are kept unchanged under the condition that the pneumatic parameters change.
The invention has the beneficial effects that:
the invention firstly establishes an air-breathing wide-speed-range aircraft dynamics model from the aspect of flight dynamics/engine thermodynamic cycle coupling. The flight/propulsion coupling characteristics of the wide-speed-range aircraft under a large-range Mach number are described by using a state space equation and a transfer function. On the basis, the flying/pushing coupling characteristics of the air-breathing wide-speed-range aircraft in the wide speed range are reproduced from the aspect of a dynamic model by building an aircraft forebody, an air inlet compression surface and a tail nozzle model. The state space equation is listed by using the disturbance linearization of the control differential equation, and an integrated coupling feedback control strategy of the wide-speed-domain aircraft under the condition of large-range Mach number is provided by adopting a pole configuration method. Compared with the traditional decoupling control strategy, the method adjusts the controller when the pneumatic parameters are changed. The control method provided by the invention obviously improves the control quality and the operation stability of the aircraft in the aspect of coping with the flying/pushing coupling effect, has the characteristics of smaller overshoot, higher response speed and fast convergence, and has a wide application prospect.
Drawings
FIG. 1 is a bin view of an aircraft airfoil;
FIG. 2 is a flow chart of aircraft inlet airflow parameter calculation;
FIG. 3 is a view of a model of the plume of the jet nozzle;
FIG. 4 is a system coupling feedback control diagram;
FIG. 5 is a flow chart of coupling control parameter modification with pneumatic parameters;
FIG. 6a is an aircraft control simulated angle of attack control curve;
FIG. 6b is an aircraft control simulated elevator control curve;
FIG. 6c is an aircraft control simulation angular rate curve;
FIG. 6d is a trajectory inclination curve simulated by aircraft control;
FIG. 6e is an aircraft control simulation pitch control curve;
FIG. 7 is a flow chart of a method for controlling the integrated propulsion coupling of a wide speed range air-breathing powered aircraft body.
Detailed Description
The following further describes a specific embodiment of the present invention with reference to the drawings and technical solutions.
The geometric modeling is carried out on the target aircraft, and the target aircraft mainly comprises a fuselage part, a wing part and an engine part. Binning is carried out on the basis of establishing a geometric model (the airfoil part of the aircraft is binned as shown in figure 1), and aerodynamic estimation is carried out by utilizing a modified Newton method and a Plantt-Meier method. Estimating the characteristics of a propulsion system of the airplane by utilizing a two-dimensional flow theory and the like to obtain the flight/thrust coupled pneumatic data and thrust data under various working conditions. And (3) linearizing the aircraft dynamics control model, and describing the flight/propulsion coupling characteristics of the wide-speed-range aircraft under a large-range Mach number by using a state space equation and a transfer function. On the basis of simplifying a state space equation, a pole allocation method is adopted, and a comprehensive control strategy of the wide-speed-domain aircraft under a large-range Mach number is given. Based on real-time changes in flight data, controller parameters are adjusted to compensate for uncertain disturbances in flight while keeping the desired pole unchanged.
The air-breathing wide-speed-range aircraft body propulsion integrated coupling effect can cause the reduction of the operation stability and the deterioration of the control quality, so the coupling feedback control is particularly important under the condition. Fig. 7 shows a method for controlling the body propulsion integration of a wide-speed-range air-breathing powered aircraft, where the embodiment takes a control task that an aircraft needs to raise an attack angle when climbing in a cruise phase as an example, and the method specifically includes the following steps:
(1) aircraft dynamics modeling
(1.1) pneumatic binning modeling
The basis of the pneumatic surface element modeling is to obtain a geometric model of the airplane, wherein the geometric model is in the form of a space coordinate data set of enough dense sampling points on the surface of the airplane. On the basis of the existing geometric model, four adjacent sampling points are used as four vertexes to form a quadrilateral surface element, and by analogy, the whole aircraft surface can be covered by the quadrilateral surface element. Recording four vertex coordinates of a certain surface element as Q 1 ,Q 2 ,Q 3 ,Q 4 And the normal vector calculation outside the surface element is as follows:
wherein, T 1 ，T 2 Is a vector obtained by subtracting two groups of non-adjacent vertexes.
The plane of a bin may be determined by the normal vector of the bin and the center of the bin. Four vertices are projected onto the surface of the bin, and the points projected onto the surface of the bin are called "corner points". The coordinates of the four corners are:
wherein
Is the center coordinates of the 4 nodes. The coordinate system is established based on a surface element, the origin is the center of the surface element, and the three axial directions are as follows: i.e. i p ＝T 1 /|T 1 |，j p ＝n×n x ，k p N. And (3) converting a calculation coordinate system into a bin coordinate system:
G cp ＝[i p ,j p ,k p ] (3)
the coordinates of the four corners in the bin coordinate system are shown in formula 4:
the coordinates of the bin centroid in the bin coordinate system are shown as formula 5:
the coordinates of the bin centroid in the calculation coordinate system are as shown in formula 6:
the area of the surface element is:
according to the engineering estimation method, the pressure on the ith surface element can be obtained.
P i ＝C pi q c,i +P ∞ (8)
Wherein P is ∞ Is the local static pressure, q c,i Is the local pressure.
The total aerodynamic and aerodynamic moments on the aircraft can be obtained by summing the aerodynamic and aerodynamic moment vectors over all surface elements.
Wherein d is i ＝d xi i b +d yi j b +d zi k b Is the distance vector from the bin centroid to the aircraft centroid, n xi ,n yi ,n zi Is the vector n at i b ,j b ,k b The component (c) above.
(1.2) supersonic pneumatic computation
After the surface element data of the whole aircraft surface is obtained, the aerodynamic force and the moment on each surface element need to be calculated for accumulation, the pressure and the pressure coefficient of each small surface element are calculated by adopting a modified Newton method (formula 13-14) and a Plantt-Meier method (formula 15), and the vector of each surface element is added, so that the total aerodynamic force and the aerodynamic moment under the working condition can be obtained. In the formula of gamma e The fuel-to-heat ratio and the airflow impact angle.
C p ＝Ksin 2 τ (13)
(1.3) Propulsion characteristic modeling
In thrust estimation, a ramjet engine is divided into an intake port, an isolator, a combustion chamber, and a nozzle. In the air inlet channel part, the flow field parameters are described by adopting shock waves, expansion waves and an intersection theory thereof. All shock and expansion waves generated in-plane due to the rotation of the inlet wall and the intersection of the shock and expansion waves were calculated. The waves divide the two-dimensional plane area into a plurality of small areas, the internal airflow parameters in each small area are the same and are determined by the shock wave and the expansion wave system in the front area. If a shock wave of shedding occurs or the number of calculated wavefronts is greater than a certain number, the average gas flow parameter will be calculated there from which the gas flow parameter to the inlet of the isolation section can be determined according to the friction variable cross-section tube flow equation (equation 16). The algorithm flow chart is shown in fig. 2.
In the formula: s. the th Is the cross-sectional area of the gas flow channel, d th Is the hydraulic diameter, where th ＝1+Ma 2 (γ c -1)/2。
And calculating airflow parameters of the isolation section and the combustion chamber by adopting a quasi-one-dimensional model, and not considering factors such as airflow friction, fuel injection, length change of a pre-combustion shock wave string, reaction rate and the like. The basic conservation equation of the airflow in the isolation section and the combustion chamber is obtained and solved by combining an experimental formula or an empirical formula. Finally, the distribution rule of the airflow parameters in the isolation section and the combustion chamber can be obtained.
Wherein: s th 、dS th Dx and d th May be determined by engine combustion chamber geometry;
and
respectively the fuel mass flow rate and the rate of change of the fuel mass flow rate, epsilon, added to the combustion chamber th Is the ratio of the fuel injection velocity to the air flow velocity on the engine shaft of the aircraft, and can be determined from the fuel injection.
The jet nozzle adopts plume modeling and reflects the coupling of flight and propulsion. The plume model is built based on the pressure equality above and below the shear layer (see fig. 3).
h k+1 ＝h k +s a tan(τ)+s a tan(β k ) (20a)
A k ＝h k+1 /h k (20b)
(2) Establishing a flight dynamics model
(2.1) establishing a wide-speed-domain flight dynamics model:
on the basis of correctly describing the motion state of the airplane and performing force analysis on the airplane, listing out an aircraft motion control differential equation.
Wherein
Is a pitch angle, alpha is an angle of attack, theta is a track inclination angle, omega z Is pitch angle velocity; the direction satisfies the right-hand rule, M z Is the pitching moment.
(2.2) linearization under the assumption of small perturbation
(2.3) introduction of simplified notation of equation coefficients
Indicating the amount of rotational velocity deviation in the ballistic ramp direction due to pitch angle velocity under current flight conditions.
Indicating that the gravity acceleration causes the deviation of the rotation angular velocity in the oblique line direction of the trajectory under the current flight condition.
The method is used for solving the problem that the angle of attack can cause the deviation of the rotation angular velocity in the oblique line direction of the trajectory under the current flight condition.
The method is used for indicating the rotating angular velocity deviation amount in the ballistic oblique line direction caused by the deflection of the control surface under the current flight condition.
The deviation of the rotation angular velocity in the ballistic oblique line direction caused by the fuel equivalence ratio under the current flight condition is shown.
And the deviation of the change of the fuel equivalence ratio caused by the attack angle under the current flight condition is shown.
a 47 ＝Kω n
The offset of the change of the fuel equivalence ratio caused by the fuel equivalence ratio command under the current flight condition is represented as the lag from the fuel equivalence ratio command to the actual fuel equivalence ratio.
(2.4) establishing a control-oriented state space equation
The corresponding transfer function is:
the transfer function order is increased relative to the transfer function of a conventional aircraft because the effect of the fuel equivalence ratio is increased.
(3) Coupled feedback control based on pole allocation
(3.1) desired pole selection: and combining the influence of the expected pole on the overshoot and the adjusting time and the actual control effect, and selecting the expected pole as follows:
j 1 ＝-3.951+4.53i
j 2 ＝-3.951-4.53i
j 3 ＝-3.435
j 4 ＝-23.825
(3.2) solving a control matrix, wherein under the selected expected pole, the control matrix is solved as follows by combining the requirements of the system dynamic response index:
(3.3) control distribution, the control matrix forms a coupled feedback control system as shown in figure 4, and the attack angle and the fuel equivalence ratio are respectively fed back to the front of the own loop controller and are also fed back to the other control loop in a cross mode. The two controllers correspond to a first row and a second row of the control matrix, respectively.
(4) Controller dynamic adjustment
The air-breathing wide-speed-range aircraft is provided with a large-envelope wire, and each pneumatic derivative value can be obviously changed under the working conditions of different heights and speeds, namely an A matrix in a state space equation is changed, so that the difficulty is brought to the control of the aircraft. In the future, the dynamic adjustment of the control matrix is needed to cope with the situation. The adjustment process is shown in fig. 5.
(4.1) when the dynamic derivative changes, first, each element C of the matrix C ═ a + BK is calculated i A value of (d);
(4.2) adding the disturbances delta A and delta B of the state matrix and the control matrix to form a new state matrix and control matrices A 'and B';
(4.3) solving the K' matrix under the condition that each element of the C matrix is kept unchanged;
(4.4) obtaining a variation value | k for each control parameter i -k i ' |, is compared with a threshold n, and if less than n, remains unchanged; if the control parameter is larger than n, correcting the actual control parameter, wherein the correction method comprises the following steps:
k i ”＝k i +λ(k i -k i ')
where λ is the correction factor.
And (4.5) calculating the change value of each control parameter, comparing, and finishing after finishing the correction.
(5) Fly/push integrated coupling controller compared with traditional controller
The flight profile of the aircraft is set at an altitude of 28000 metres and at a speed of mach 6.5. The initial angle of attack is 0 deg. and is adjusted to 1 deg.. The initial fuel equivalence ratio was 0, adjusted to 0.85, and simulated in 0 to 10 seconds. The simulation result comprises a dynamic response process under a normal working condition, an elevator dynamic response process after a certain amount of uncertainty is added, and a dynamic response process of controller self-adaptive compensation when the uncertainty is added.
TABLE 1 simulation parameters
For the anti-disturbance control strategy of the airplane in the wide speed range, the established model and method are utilized to design and simulate the terminal flight speed of Ma6.5, the terminal height, the terminal flight trajectory angle and the flight attack angle adjusted from 0 degree to 0.5 degree. The elevator dynamic response process after accounting for 20% uncertainty and the dynamic response process with controller adaptive compensation when adding 20% uncertainty. The simulation results are shown in fig. 6 a-6 e. The important system performance index information is summarized in the following table:
TABLE 2 dynamic Performance index comparison
As can be seen from fig. 6a and 6b, the maximum overshoot of the attack angle by the control system is 3.7%, and the adjustment time is 2.17 seconds, which substantially meets the requirement of the expected endpoint on the dynamic response index of the system. Compared with the traditional flight control method, the maximum overshoot is 11.6%, the adjusting time is 4.13 seconds, and simulation results show that the flight/push integrated coupling control method can obviously improve the control quality of the uncertain air-breathing wide-speed-range aircraft system. As can be seen from fig. 6 c-6 e, the integrated fly/push coupling control can reduce the maximum peak of the curve. As can be seen from fig. 6c and 6d, the response curve of the fly/push integrated coupling control is closer to the steady state value as a whole.
Claims (1)
1. The integrated coupling control method for the propulsion of the wide-speed-range air-breathing power aircraft body is characterized by comprising the following specific steps of:
(1) aircraft dynamics modeling
(1.1) pneumatic binning modeling
Dividing the surface of the aircraft into a plurality of small areas by using a surface element method, replacing the small areas with parallelograms or triangles, and calculating aerodynamic force and moment on a surface element;
(1.2) supersonic pneumatic computation
Calculating the pressure and pressure coefficient of each surface element by adopting an ultrasonic pneumatic estimation method, and adding the vectors of each surface element to obtain the total aerodynamic force and the aerodynamic moment under the working condition;
(1.3) propulsion System calculations
In the thrust estimation, the scramjet engine is divided into an air inlet channel, an isolator, a combustion chamber and a spray pipe;
in the air inlet part, describing flow field parameters by adopting shock waves, expansion waves and an intersection theory thereof, and calculating all shock waves and expansion waves generated in a plane due to the rotation of an inlet wall and the intersection of the shock waves and the expansion waves; the waves divide a two-dimensional plane area into a plurality of small areas, the internal airflow parameters in each small area are the same and are determined by shock waves or expansion waves in the front area;
calculating airflow parameters of the isolator and the combustion chamber by adopting a quasi-one-dimensional model, obtaining a basic conservation equation of the airflow in the isolator and the combustion chamber without considering airflow friction, fuel injection, length change of a pre-combustion shock wave string and reaction rate, and solving by combining an empirical formula;
the plume modeling is adopted for the tail nozzle, a two-dimensional model of the tail nozzle is divided into small areas along a body axis, and the plume model is built on the basis of equal air pressure above and below the shear layer;
(2) analysis of dynamics characteristics of wide-speed-range air-breathing dynamic aircraft
(2.1) establishing a wide-speed-domain flight dynamics model:
in the formula: m is the pitching moment, omega z Is the pitch angular velocity, J z Is the pitch moment of inertia, V is the velocity, theta is the track inclination, H is the flight altitude, m is the aircraft mass,
p, X, Y respectively represents thrust, drag and lift of the engine, g represents gravity acceleration, and alpha represents an attack angle;
due to the obvious fly-push coupling characteristic of the wide-speed-range air-breathing power aircraft, description of an actual fuel equivalence ratio needs to be added, the fuel equivalence ratio is related to speed, height and a fuel equivalence ratio instruction, and an expression is as follows:
wherein the content of the first and second substances,
refers to the command of the fuel equivalence ratio,
is a fuel equivalence ratio command;
(2.2) carrying out linearization expansion on the aircraft in a certain flight state according to a small disturbance linearization theory; the expansion form is that each line of the equation is formed
In the form of (a);
(2.3) in order to make the expression after the flight dynamics formula linearization easy to write, introducing a simplified symbol of an equation coefficient; the static stability and the static maneuverability of the airplane are important reference indexes for evaluating the stability of the airplane, and the specific indexes are as follows:
a 24 the static stability of the airplane is represented, and is the acceleration deflection of the rotation angle of the airplane around the body axis caused by the change of a unit attack angle; wherein the content of the first and second substances,
for changes in pitch moment due to aerodynamic forces resulting from changes in unit angle of attack,
change in pitching moment due to thrust resulting from change in unit angle of attack, J z Is the moment of inertia;
a 25 the characteristic of static maneuverability, in particular the efficiency of elevators, is the O-turn of the aircraft caused by the unit deflection of the steering mechanism z1 Shaft rotation angle acceleration deflection; wherein the content of the first and second substances,
is caused by unit deflection of steering mechanism z1 The moment offset of the shaft is determined by the amount of moment offset,
is caused by unit deflection of steering mechanism z1 Moment coefficient offset of the shaft, q is the dynamic pressure of the oncoming airflow, S is the reference area of the aircraft, b A Is the average aerodynamic chord length;
(2.4) establishing a control-oriented aircraft/propulsion integrated coupling state space equation:
from the aircraft/propulsion integrated coupling state space equation, the system has two control quantities which are respectively an elevator and a fuel equivalence ratio instruction; the state quantities are four, namely fuel equivalence ratio, angular velocity, pitch angle and attack angle;
(3) coupled feedback control based on pole allocation
Coupling feedback control is realized through pole allocation of a multi-input multi-output system, the control process comprises the steps of selecting an expected pole position according to expected overshoot and adjusting time, solving controller parameters and controlling distribution, and the specific process comprises the following steps:
(3.1) desired Pole selection
According to longitudinal short-period disturbance motion
The system state space equation after the control signal is added is obtained from the time domain perspective as follows:
if the value of the K matrix is such that A + BK forms a matrix that stabilizes the system asymptotically, with x (0) ≠ 0, then
The eigenvalues of A + BK are the poles of the regulator; if the pole of the regulator is located in the left half-plane of the coordinate system, then
The problem of placing the poles of the regulator in the desired positions is called pole placement;
the process of desired pole selection is as follows:
according to the logic of the elevator control attitude and the fuel equivalence ratio control system, the fourth-order system is decomposed into a third-order system and a first-order system which respectively correspond to the attack angle control system and the fuel equivalence ratio control system; the three-order system is provided with two dominant poles and a non-dominant pole, and the dynamic index of the step response of the system is mainly determined by the two dominant poles; selecting a dominant pole by calculating a damping ratio and a natural frequency; the formula is as follows:
wherein s is 1,2 Representing two dominant poles, ξ being the system damping ratio, ω n Is the natural frequency of the system, t s To adjust time, σ is the system overshoot;
after the selection of the two dominant poles is completed, according to the real part of the dominant pole, the third pole, i.e. the non-dominant pole, is autonomously selected, and a value more than 5 times that of the dominant pole is usually selected as the non-dominant pole, i.e.:
s 3 ＝-nξω n
n is a selected multiple and is adjusted through an actual control effect;
the first order system has one pole, noted:
s 4 ＝-1/T
wherein T is a first-order system time constant, and a pole of the first-order system is selected by adjusting time;
to this end, the positions of the four desired poles are all determined:
(3.2) control parameter solving
(3.3) control of dispensing
The logic of the control matrix is:
when calculating the adjustment quantity delta of the elevator z In the process of (a), k 11 Δω z And k 12 Δ α is a differential term and a proportional term, respectively;
is a coupled control item of the flight/propulsion system,
the expressed fuel equivalence ratio error integral term is independent of elevator adjustment, so k 14 0; similarly, the fuel equivalence ratio command adjustment is solved
In the course of (a) or (b),
and
respectively representing a proportional term and an integral term, k 22 Δ α is a coupling term, k 21 Δω z Is an unrelated item;
(4) controller dynamic adjustment
In the wide-speed-range air-breathing power aircraft, pneumatic parameters may significantly change in the wide-speed-range flight process, and the control matrix K is adjusted through cyclic detection, so that the element values of the A + BK matrix are kept unchanged under the condition that the pneumatic parameters change.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115981161A (en) * | 2023-02-07 | 2023-04-18 | 大连理工大学 | Fixed-wing aircraft gliding control gain index optimization method based on direct lift force |
Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103970957A (en) * | 2014-05-20 | 2014-08-06 | 北京航空航天大学 | Simulation method for elastic waverider hypersonic flight vehicle |
| CN107967382A (en) * | 2017-11-20 | 2018-04-27 | 中国运载火箭技术研究院 | A kind of airbreather coupling influence degree evaluation method |
| CN107977009A (en) * | 2017-11-20 | 2018-05-01 | 中国运载火箭技术研究院 | A kind of airbreather attitude control law design method for considering coupling |
| CN108052008A (en) * | 2018-01-10 | 2018-05-18 | 大连理工大学 | A kind of geometry controller design method of tiltrotor aircraft transition state handoff procedure |
| CN109101765A (en) * | 2018-09-19 | 2018-12-28 | 西北工业大学 | A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft |
| CN110187715A (en) * | 2019-04-24 | 2019-08-30 | 南京航空航天大学 | A kind of hypersonic aircraft control method for coordinating based on dynamic Coupling Analysis |
| CN110287505A (en) * | 2019-03-20 | 2019-09-27 | 北京机电工程研究所 | Stability of aircraft analysis method |
| CN110908278A (en) * | 2019-11-12 | 2020-03-24 | 北京航空航天大学 | Dynamics modeling and stability control method of folding wing aircraft |
| CN111125971A (en) * | 2019-12-26 | 2020-05-08 | 北京航空航天大学 | Method for determining thrust uncertainty of air-breathing hypersonic aircraft |
| CN111274740A (en) * | 2020-01-10 | 2020-06-12 | 中国人民解放军国防科技大学 | Multi-aircraft cooperative penetration trajectory optimization design method |
| CN114313253A (en) * | 2022-03-03 | 2022-04-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Aerodynamic layout and design method of high lift-drag ratio air-breathing hypersonic aircraft |
-
2022
- 2022-05-24 CN CN202210566517.1A patent/CN114912202A/en active Pending
Patent Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103970957A (en) * | 2014-05-20 | 2014-08-06 | 北京航空航天大学 | Simulation method for elastic waverider hypersonic flight vehicle |
| CN107967382A (en) * | 2017-11-20 | 2018-04-27 | 中国运载火箭技术研究院 | A kind of airbreather coupling influence degree evaluation method |
| CN107977009A (en) * | 2017-11-20 | 2018-05-01 | 中国运载火箭技术研究院 | A kind of airbreather attitude control law design method for considering coupling |
| CN108052008A (en) * | 2018-01-10 | 2018-05-18 | 大连理工大学 | A kind of geometry controller design method of tiltrotor aircraft transition state handoff procedure |
| CN109101765A (en) * | 2018-09-19 | 2018-12-28 | 西北工业大学 | A kind of wide fast domain propulsion system modelling by mechanism method of big envelope curve of assembly power aircraft |
| CN110287505A (en) * | 2019-03-20 | 2019-09-27 | 北京机电工程研究所 | Stability of aircraft analysis method |
| CN110187715A (en) * | 2019-04-24 | 2019-08-30 | 南京航空航天大学 | A kind of hypersonic aircraft control method for coordinating based on dynamic Coupling Analysis |
| CN110908278A (en) * | 2019-11-12 | 2020-03-24 | 北京航空航天大学 | Dynamics modeling and stability control method of folding wing aircraft |
| CN111125971A (en) * | 2019-12-26 | 2020-05-08 | 北京航空航天大学 | Method for determining thrust uncertainty of air-breathing hypersonic aircraft |
| CN111274740A (en) * | 2020-01-10 | 2020-06-12 | 中国人民解放军国防科技大学 | Multi-aircraft cooperative penetration trajectory optimization design method |
| CN114313253A (en) * | 2022-03-03 | 2022-04-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Aerodynamic layout and design method of high lift-drag ratio air-breathing hypersonic aircraft |
Non-Patent Citations (2)
| Title |
|---|
| JIAXIN LI, KAI LIU, ET AL.: "An_Integrated_Control_Approach_for_Air-breathing_Hypersonic_Vehicle" * |
| JIAXIN LI,DANWEI LI,ET AL.: "Flight-Propulsion_Integration_Dynamic_Analysis_and_Adaptive_Control_of_the_Hypersonic_Vehicle_at_Wide-Range_Mach_Numbers" * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115981161A (en) * | 2023-02-07 | 2023-04-18 | 大连理工大学 | Fixed-wing aircraft gliding control gain index optimization method based on direct lift force |
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