CN116301028A - Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform - Google Patents
Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform Download PDFInfo
- Publication number
- CN116301028A CN116301028A CN202310088829.0A CN202310088829A CN116301028A CN 116301028 A CN116301028 A CN 116301028A CN 202310088829 A CN202310088829 A CN 202310088829A CN 116301028 A CN116301028 A CN 116301028A
- Authority
- CN
- China
- Prior art keywords
- flight
- angle
- track
- section
- attack
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 103
- 238000013461 design Methods 0.000 claims abstract description 24
- 239000000446 fuel Substances 0.000 claims abstract description 17
- 230000008878 coupling Effects 0.000 claims abstract description 5
- 238000010168 coupling process Methods 0.000 claims abstract description 5
- 238000005859 coupling reaction Methods 0.000 claims abstract description 5
- 230000008569 process Effects 0.000 claims description 24
- 230000009194 climbing Effects 0.000 claims description 21
- 230000008859 change Effects 0.000 claims description 16
- 239000011159 matrix material Substances 0.000 claims description 16
- 230000001174 ascending effect Effects 0.000 claims description 7
- 230000001133 acceleration Effects 0.000 claims description 6
- 238000013016 damping Methods 0.000 claims description 6
- 230000005484 gravity Effects 0.000 claims description 6
- 230000001105 regulatory effect Effects 0.000 claims description 6
- 230000000630 rising effect Effects 0.000 claims description 5
- 238000006243 chemical reaction Methods 0.000 claims description 4
- 230000010354 integration Effects 0.000 claims description 4
- 230000007704 transition Effects 0.000 claims description 4
- 230000005540 biological transmission Effects 0.000 claims description 2
- 125000004122 cyclic group Chemical group 0.000 claims description 2
- 230000001419 dependent effect Effects 0.000 claims description 2
- 238000005457 optimization Methods 0.000 claims description 2
- 230000002028 premature Effects 0.000 claims description 2
- 230000004044 response Effects 0.000 claims description 2
- 239000007787 solid Substances 0.000 claims description 2
- 230000008901 benefit Effects 0.000 abstract description 6
- 238000004364 calculation method Methods 0.000 description 4
- 238000010586 diagram Methods 0.000 description 3
- 238000011160 research Methods 0.000 description 3
- 238000004088 simulation Methods 0.000 description 3
- 238000013459 approach Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 239000010763 heavy fuel oil Substances 0.000 description 2
- 239000003921 oil Substances 0.000 description 2
- 239000003380 propellant Substances 0.000 description 2
- 238000013528 artificial neural network Methods 0.000 description 1
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000004422 calculation algorithm Methods 0.000 description 1
- 230000001276 controlling effect Effects 0.000 description 1
- 238000011156 evaluation Methods 0.000 description 1
- 238000012423 maintenance Methods 0.000 description 1
- 238000003062 neural network model Methods 0.000 description 1
- 239000007800 oxidant agent Substances 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
- 229910052760 oxygen Inorganic materials 0.000 description 1
- 239000001301 oxygen Substances 0.000 description 1
- JTJMJGYZQZDUJJ-UHFFFAOYSA-N phencyclidine Chemical compound C1CCCCN1C1(C=2C=CC=CC=2)CCCCC1 JTJMJGYZQZDUJJ-UHFFFAOYSA-N 0.000 description 1
- 239000002760 rocket fuel Substances 0.000 description 1
- 230000000007 visual effect Effects 0.000 description 1
Images
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention belongs to the technical field of hypersonic aircraft middle section guiding, and relates to a multi-constraint online flight track planning middle section guiding method based on an air suction hypersonic platform. The invention firstly designs the flight track on line based on the air-pushing coupling characteristic of the air suction type platform. In the transverse plane, the transverse plane turns based on a fixed roll angle, and the track design is performed by eliminating the lead angle. In the longitudinal plane, a principle that the fuel consumption is low and the dynamic pressure state can be kept for a long time is adopted to design a longitudinal flight strategy. The method has great advantages in oil consumption, and can effectively enlarge the area of an interception airspace. After the nominal track is obtained, track section tracking guidance is realized by adopting a track linearization control method, and the effectiveness of the method is verified. The method is suitable for middle section guidance based on the suction type super platform and has wide application prospect.
Description
Technical Field
The invention belongs to the technical field of hypersonic aircraft middle section guiding, and relates to a multi-constraint online flight track planning middle section guiding method based on an air suction hypersonic platform.
Background
The middle section guiding task is to guide the suction hypersonic platform to the working distance range of the infrared guide head. In recent years, various mid-section guidance schemes have been developed for conventional hypersonic aircraft. The sliding mode guidance, the proportional guidance and the optimal control theory are more and more widely applied to the design of the middle-section guidance law.
In the published research, the guidance law in the middle section of the sliding mode can enable the guidance system to converge to the sliding mode surface in a limited time, so that the sight angle is ensured to converge to an expected value when the middle terminal guidance is shifted, and the sight angle speed is close to zero. On the basis, the interception performance requirement is considered, and a quadratic optimal control method is adopted to solve the middle-section guidance comprising constraint conditions and optimal control problems. The guidance law has the advantage of controlling energy optimally, but also has the problems of mathematical complexity and large calculation amount. It is not acceptable to consider complex flight calculations. Therefore, the guidance system provides higher requirements for the design method of the guidance law of the middle section. Still other studies use neural networks to approximate optimal feedback strategies suitable for real-time implementation. The suboptimal middle guidance law is obtained through the offline trained neural network model, and the method has a good approximation effect on the optimal flight trajectory. The algorithm has the characteristics of high calculation efficiency, high precision and suitability for airborne guidance.
The above research method and research content are mainly applied to the traditional hypersonic aircraft. The scramjet engine adopted by the air suction hypersonic speed platform can extract oxygen from the atmosphere in the flight process, does not need to carry an oxidant, and reduces the load quality. This means that for the same mass of propellant, a scramjet can produce more thrust than a rocket engine. In addition, scramjet engines have severe dynamic pressure constraints, and overrun may cause the engine to stall, resulting in intercept failure. However, the existing middle-stage guiding design method cannot effectively utilize the working characteristics of the scramjet engine, and brings difficulties to the design of the middle-stage guiding method based on the air suction type platform.
Disclosure of Invention
Because the scramjet engine equipped on the suction hypersonic speed platform has special requirements on the running condition, the dynamic pressure is required to strictly meet the constraint. In addition, considering that the difference of fuel consumption modes between rocket power and a scramjet engine causes inapplicability of the traditional guidance method, the problem of long calculation time exists in the process of remote interception of the optimal middle section guidance. Based on the method, the method is designed for redesigning the online track planning method aiming at the dynamic characteristics of the air suction hypersonic speed platform, the multi-constraint middle section guiding strategy method based on the air suction hypersonic speed platform is designed, a new prediction hit point is planned online based on a residual time estimation method, the platform characteristics are combined, the flight track meeting the state constraint is designed online based on a relative position model, online guidance is completed according to track tracking control, and online track planning and guidance are completed rapidly and accurately. Thereby providing a related technical approach for the design of the middle guidance law of the suction type hypersonic platform.
The technical scheme of the invention is as follows:
a multi-constraint on-line flight path planning middle section guiding method based on an air suction hypersonic speed platform comprises the following steps:
(1) Online flight trajectory planning method based on predicted hit point
(1.1) air-breathing hypersonic platform flight strategy
After the air suction type hypersonic platform is launched, the air suction type hypersonic platform is subjected to two flight states of a boosting section and a scramjet power cruising section.
(a) Boosting section: the boosting section is performed by adopting a rising section instruction generation rule. According to the method, the ascending process of the aircraft is divided into a plurality of flight stages, fixed program angle instructions are set in each stage, the instructions are determined by a limited number of parameters, the parameters and corresponding program instructions are obtained through offline optimization according to task requirements, the parameters and the corresponding program instructions are bound in a booster system, and the program instructions are realized by using a gesture control system in the ascending stage of the aircraft. To seek rapid generation of flight trajectory, the entire rising section is divided into 3 sections including a vertical rising section (T is more than or equal to 0 and less than or equal to T) 11 ) Negative angle of attack turning section (T) 11 ≤t≤T 12 ) Gravity turning section (T) 12 ≤t≤T 13 )。
The whole-course flight attack angle approximation formula of the suction hypersonic speed platform is as follows:
wherein: 1/v 0 Is the thrust-weight ratio. Alpha m Is the maximum negative angle of attack magnitude, t m At the maximum moment of negative attack angle value, T 12 The negative attack angle turning section can be freely arranged; t (T) 13 Is the solid rocket fuel exhaustion time.
(b) Scramjet power cruise section: for an air suction hypersonic platform, the thrust provided by the scramjet engine can be obtained through data interpolation:
wherein Q is dynamic pressure; ρ is the air density at the height of the aspirated hypersonic platform; k (k) r The opening of the valve is regulated for thrust; c is the sound velocity; ma is Mach number; t is the engine thrust. V is the aspiration hypersonic plateau velocity.
(b.1) valve opening: since the cruising speed is greater than the speed at the end of the boost segment, this requires that the aspirated hypersonic platform first accelerate to a specified speed and then turn into a constant speed flight. The thrust is the maximum value provided by the scramjet engine in the accelerating flight process, namely the valve opening k r =1, i.e.:
T＝T max
in order to maintain a constant speed in a constant speed flight state, it is necessary to make
Namely:
wherein: θ, δ, m, r, x, y, z, α represent three components of the ballistics dip angle, ballistics deflection angle, mass, centroid distance and emission system of the aspiration hypersonic platform, respectively, and the attack angle; t, D the engine thrust and drag, respectively; g is the gravitational acceleration of the earth surface.
The opening of the valve is as follows:
(b.2) angle of attack: the cruise section considers the flight characteristics of the suction hypersonic platform and requires the hypersonic platform to be in a flight state with equal high and constant speed as much as possible. To meet the intercept demand, the flight trajectory may be divided into a flat flight segment and a climb segment based on the altitude of the flight. The angle of attack instruction is obtained from program instructions for different flight conditions.
(b.2.1) Flat fly section
The flat flight section requires the change rate of the altitude of the suction hypersonic platform to be unchanged, so that an aircraft is required
As a result of:
h＝r-R e
wherein: h. re represents the hypersonic platform fly height and the earth radius, respectively.
Can derive x, y and z and further calculate three components V of the speed of the transmission system x 、V y 、V z Deriving to obtain the second derivative change rate of the attack angle:
due to
Related to rise and resistance, which in turn are related to angle of attack, by setting an initial iteration value alpha of angle of attack 0 Small angle of attack Δα by iteration +.>
The angle of attack is obtained by means of (a).
Wherein: alpha k Is the attack angle value after iterating k times, lambda is the optimal step factor of the damping Newton method, and can enable:
wherein: d, d k Is the search direction, order
Namely:
wherein:
wherein:
is the attack angle alpha k Acceleration of each axis of the emission coordinate system at the moment, +.>
Is the attack angle alpha k Acceleration of each axis of emission coordinate system at +Δα, setting cutoff condition |α k+1 -α k |<And epsilon are small, and iterating until the cut-off condition is met, so that the attack angle value of the plane flight segment can be obtained.
(b.2.2) climbing section
The climbing section refers to a flight process of the hypersonic platform climbing to a designated height, and can be subdivided into three parts of a constant attack angle climbing section, a linear climbing section and a smooth transition section.
Firstly, the hypersonic platform changes the inclination angle of a flight track by increasing lift force through a large attack angle in a fixed attack angle climbing section, so that the hypersonic platform has climbing capability.
When the inclination angle of the flight track reaches a preset value, the flight track ascends to a certain height in a linear climbing mode. In order to keep the intercepted flight path to climb straight, the inclination angle change rate of the flight path is required to be constant at 0 degrees/s, and the inclination angle change rate of the flight path is as follows:
iterative initial value alpha of attack angle can be set 0 By newton's iteration, namely:
by cyclic iteration of angle of attack alpha k To
And obtaining a flight attack angle alpha as an iteration termination condition, and obtaining a linear climbing flight track by integration.
And finally, obtaining a program instruction of the flight track inclination angle through a smooth transition section by taking the flight height as an independent variable and taking the flight track inclination angle as an excessive function of the dependent variable.
θ d (h)＝φ(h)*θ 1 +(1-φ(h))θ 0
Wherein: h is a max To intercept the height, h min For a given smooth segment start height, h is the actual height, θ 0 To smooth the initial flight path inclination angle theta 1 To expect the inclination angle of the flight track, the inclination angle theta of the flight track is expected under the coordinate system of the flight track at the smooth tail end due to the design requirement of equal height and constant speed 1 ＝0°The flight trajectory tilt command can be reduced to:
θ d (h)＝(1-φ(h))θ 0
therefore, in order to convert the flight trajectory inclination angle instruction in the flight trajectory coordinate system into the flight trajectory inclination angle instruction in the emission system, it is necessary to obtain by conversion through the coordinate conversion matrix:
wherein: delta d As any one standard flight track group exists only in the longitudinal plane, so that v=delta d Identical to 0 deg.. After the flight track inclination angle theta' under the emission coordinate system is obtained, the change rate of the flight track inclination angle can be obtained based on the iteration step delta t by subtracting the flight track inclination angle theta of the current actual emission coordinate system from the flight track inclination angle theta of the current actual emission coordinate system:
the rate of change of the inclination of the flight trajectory in the emission coordinate system can also be expressed as:
and (3) finishing iteration of the attack angle through the change rate of the inclination angle of the flight track, substituting the iteration into the dynamic model integration to obtain the flight track of the smooth section, wherein the specific method is the same as that of the flat flight section.
(1.2) design of on-line flight trajectory planning method
Based on the air-breathing platform air-pushing coupling characteristic, the flight track is designed on line, wherein the control parameters are attack angle, tilting angle and engine throttle.
In the transverse plane, consider yaw path, yaw path turns based on fixed roll angle, track design by eliminating the lead angle, and give roll angle instruction based on the direction of lead angle:
wherein eta zmin To allow maximum deviation of line of sight angle, gamma 0 For a set fixed roll angle magnitude.
In the longitudinal plane, the problem that the optimal working state of the scramjet engine adopted by the suction hypersonic speed platform is an equal dynamic pressure flight state and the influence of the flight height on the fuel consumption rate is large is considered. Here, longitudinal flight strategies are designed based on the principle of low fuel consumption and maintenance of an equivalent dynamic pressure state over a long period of time.
It is known that dynamic pressure is related to flying speed and altitude, and for an aspiration hypersonic platform flying at a constant speed, dynamic pressure is related to flying altitude only, and thus an equal dynamic pressure state can be maintained by maintaining altitude unchanged. In terms of fuel consumption rate, considering that the influence of the impact of flying height contrast is larger, the fuel consumption can be effectively reduced by adopting a low flying track, so that the subsequent interception is more facilitated. Based on the above principle, here based on the predicted hit point height h f The current moment height h of hypersonic speed platform 0 The flight trajectory is designed based on the altitude deviation.
(a) If the current flight height h is larger than the predicted hit point height, the characteristic of low flight track specific impulse is utilized, firstly, the downward flight track is adopted, the height of the predicted hit point is reduced, and then the cruise flight mode is adopted to fly against the predicted hit point.
(b) If the current flight height h is smaller than the predicted hit point height, in order to avoid premature climbing, specific impact is reduced, fuel consumption is uneconomical, the aircraft flies at the current flight height, climbs to the interception height after the remaining flight time is smaller than a preset value, and returns to the cruise flight mode again.
(2) Nominal track tracking guidance method design
The nominal trajectory is obtained by step (1). Track following guidance can be described as designing a proper guidance law after designing a nominal track according to an on-line flight track planning method based on a predicted hit point, so that the actual flight track of the middle guide section well tracks the nominal track.
The trajectory dynamics model is dimensionally expanded by taking the speed, the altitude and the transverse position as state variables, and the state space is expressed as
Wherein x (t) is the state quantity of the middle guide section in the flight process; u (t) is a control variable.
Assume that
In the method, in the process of the invention,
is a nominal track state quantity; />
Is a nominal trajectory control quantity; e is the difference between the actual flight trajectory and the nominal trajectory state quantity, +.>
Is the difference between the control quantity and the nominal track control quantity in the actual flight process. Then there is
Along e (t) =0,
linearization, can obtain
In the method, in the process of the invention,
for the above system, a time-varying controller is used, which takes the form:
in the method, in the process of the invention,
where K (t) is a 3X 6 matrix of control parameters. Substituting the above formula into the time-varying controller can obtain a closed-loop system matrix.
Let the desired closed-loop matrix be
Wherein lambda is i (i=1, …, 6) is the desired feature root.
The closed-loop system matrix is equal to the closed-loop system matrix obtained above, and each control parameter in the expected closed-loop matrix can be obtained
In the method, in the process of the invention,
the control parameters are determined by the desired feature root, so that the desired feature root needs to be selected appropriately to obtain the control parameters.
The three subspaces can be described as a second order system, described in terms of a characteristic equation, as follows:
the solution can be obtained:
in the method, in the process of the invention,
is natural frequency omega 1 (t),ω 2 (t),ω 3 (t) damping ratio, by demand performance to determine the desired system +.>
And omega 1 (t),ω 2 (t),ω 3 (t) thereby designing controller parameters that meet performance requirements.
For an underdamped second order linear system, the step response rise time and overshoot estimation formulas are as follows:
the rising time and overshoot of the system can be obtained by the above formula, and meanwhile, the damping ratio and the frequency of the expected system can be obtained, so that the characteristic root of the expected system matrix can be determined, and the controller parameters can be obtained.
The invention has the beneficial effects that:
according to the invention, firstly, a flight track needs to be designed on line based on the air-pushing coupling characteristic of the air-sucking platform, wherein control parameters are an attack angle, a roll angle and an engine throttle. In the yaw path, the yaw path makes a turn based on a fixed roll angle, performs a trajectory design by eliminating a lead angle, and gives a roll angle command based on the direction of the lead angle. In the longitudinal plane, a principle that the fuel consumption is low and the dynamic pressure state can be kept for a long time is adopted to design a longitudinal flight strategy. Compared with the traditional middle section guiding method of proportional guiding, the on-line flight path planning middle section guiding method limits the flight height of the platform, thereby ensuring that dynamic pressure meets the working condition of the scramjet engine. In the aspect of fuel consumption, the fuel-saving track meeting the condition of equal high constant speed is planned in the middle section of the online flight track planning. On the basis, a relative position model is combined, and a guidance instruction for ensuring that the hypersonic speed platform is positioned on the most fuel-saving track most of the time is designed, so that the fuel consumption is reduced. Therefore, the guiding in the middle section of the online flight path planning has great advantage in the aspect of oil consumption, and the area of an interception airspace can be effectively enlarged. The method is suitable for middle section guidance based on the suction type super platform and has wide application prospect.
Drawings
FIG. 1 is a flow chart of a multi-constraint on-line flight trajectory planning mid-section guidance method based on an aspiration hypersonic platform;
FIG. 2 is a schematic diagram of trajectory planning for a longitudinal flight strategy;
FIG. 3 is a mid-section guidance map of online trajectory planning;
FIG. 4 is an angle of attack versus time graph of a trajectory online planning method;
FIG. 5 is a graph of track roll angle versus time for a track on-line planning method;
FIG. 6 is a graph of thrust regulating valve opening versus time for a trajectory online planning approach;
FIG. 7 is a graph of gravity, lift versus time for a trajectory on-line planning method;
FIG. 8 is a graph of thrust, drag versus time for a trajectory on-line planning method;
FIG. 9 is a height-time plot of a trajectory online planning method;
FIG. 10 is a specific impulse versus time curve for a trajectory online planning method;
FIG. 11 is a mass versus time plot of a trajectory online planning method;
FIG. 12 is an angle of attack versus time graph for track linearization guidance control;
FIG. 13 is a roll angle versus time graph of a track linearized guidance control;
FIG. 14 is a throttle-time graph of a track linearized guidance control.
Detailed Description
The following describes the embodiments of the present invention further with reference to the drawings and technical schemes.
A multi-constraint on-line flight trajectory planning middle-stage guiding method based on an air suction hypersonic platform comprises an on-line flight trajectory planning method based on a predicted hit point and a nominal trajectory tracking guiding method design. The flow chart of the middle guide method is shown in fig. 1.
The embodiment is specifically as follows:
(1) Inputting an initial state, giving a target state
Assuming that the initial position of the interception array is (-3.9 degrees, 79.8 degrees), the initial predicted hit point of the suction hypersonic platform is (3.19 degrees, 79.18 degrees, 29 Km), the initial emission azimuth angle is-5 degrees, the flying speed is 1800m/s, the cruise+climb+cruise flying mode is adopted, after 105s, 350s and 400s are emitted, the predicted flight track is refreshed, the predicted hit point is changed to (3.00 degrees, 80.00 degrees, 28 Km), (2.90 degrees, 79.90 degrees, 29 Km), (2.95 degrees, 79.95 degrees, 29.5 Km), and the position change straight line distance is respectively larger than 100Km,10Km and 5Km.
(2) On-line flight path planning method design based on predicted hit point
After the predicted hit point is obtained, the flight trajectory needs to be designed online based on the air-push coupling characteristic of the air-suction platform. In the yaw path, the yaw path makes a turn based on a fixed roll angle, performs a trajectory design by eliminating a lead angle, and gives a roll angle command based on the direction of the lead angle. In the longitudinal plane, a principle that the fuel consumption is low and the equal dynamic pressure state can be maintained for a long time is adopted to design a longitudinal flight strategy, and a track planning schematic diagram of the longitudinal flight strategy is shown in figure 2. And comparing the indexes such as the position error of the terminal, the line-of-sight angular rate, whether the residual fuel and the dynamic pressure exceed the state constraint and the like with the proportional guidance, and verifying the superiority of the middle-section guidance method based on the online flight path planning.
Firstly, performing performance analysis on the middle guide based on the proportional guide by utilizing simulation conditions of online track planning simulation. The guiding schematic diagram of the middle section of the online track planning is shown in fig. 3; as can be seen from FIG. 3, the course can be corrected within a specified time by adopting the on-line track planning middle-section guiding method, and the predicted hit point is reached, so that the method has a certain use value.
FIG. 4 is an angle of attack versus time graph of a trajectory online planning method; FIG. 5 is a graph of track roll angle versus time for a track on-line planning method; FIG. 6 is a graph of thrust regulating valve opening versus time for a track on-line programming method; FIG. 7 is a graph of gravity, lift versus time for a trajectory on-line planning method; FIG. 8 is a graph of thrust, drag versus time for a trajectory on-line planning method; from the above graph, in the whole flight process, the attack angle change is gentle, the opening of the thrust regulating valve floats in a reasonable range, the lift force and the gravity can reach balance after the predicted hit point is changed, the thrust force and the resistance can also reach balance after the predicted hit point is changed, and the aircraft can fly towards the target cruising after the predicted hit point is changed. And smoothly guide to the target
To further demonstrate the performance superiority of the on-line flight trajectory planning mid-section guidance method, fuel consumption and dynamic pressure were compared. As can be seen from the height-time curve of fig. 9, the in-line flight trajectory planning mid-section guidance limits the flight height of the platform, thereby ensuring that the dynamic pressure meets the conditions of the scramjet engine. From the specific impulse-time curve of fig. 10, it can be seen that the specific impulse of the aircraft is maintained at a high level during the flight, which means that the efficiency of the engine is high, and the speed increment that can be generated by the propellant under the same conditions is larger, which means that the method can fully exert the advantages of the scramjet engine of the air suction type platform.
In the aspect of fuel consumption, the fuel-saving track meeting the condition of equal high constant speed is planned in the middle section of the online flight track planning. On the basis, a relative position model is combined, and a guidance instruction for ensuring that the hypersonic speed platform is positioned on the most fuel-saving track most of the time is designed, so that the fuel consumption is reduced. The mass-time curve is shown in fig. 11. Therefore, the guiding in the middle section of the online flight path planning has great advantage in the aspect of oil consumption, and the area of an interception airspace can be effectively enlarged. Table 1 is a comparison of the on-line flight trajectory planning mid-segment guidance method and the proportional guidance method in terms of terminal error, line of sight angular rate, residual fuel, and dynamic pressure constraints.
Table 1 data comparison table
As can be seen from Table 1, the proportional guidance has the problem of overlarge line-of-sight angular rotation rate, and the on-line flight trajectory planning middle-stage guidance method can ensure that the line-of-sight angular rotation rate is eliminated to zero, thereby providing better interception conditions for the subsequent interception. In conclusion, all evaluation index performances of the on-line flight path planning middle-section guidance are better than those of the proportional guidance, and the on-line flight path planning middle-section guidance can effectively improve the interception performance of the platform.
(3) Nominal track tracking guidance method design
After the nominal track is obtained, in order to verify the effectiveness of the track linearization guidance law, guidance simulation analysis considering the lift pull deflection by 10% is performed, and the effectiveness of the method is verified. The usable attack angle range alpha epsilon minus 4 degrees, 6 degrees of the aircraft, and the usable roll angle range gamma epsilon minus 50 degrees, 50 degrees of the aircraft.
As can be seen from the attack angle-time curve chart of FIG. 12, the roll angle-time curve chart of FIG. 13 and the thrust regulating valve opening-time curve chart of FIG. 14, the track linearization guidance control method of the invention has the advantages that the attack angle, the roll angle and the throttle command achieve good tracking effect, the control parameter change is mild, the anti-interference capability is strong, and the control load of the aircraft can be effectively reduced. The track linear guidance has higher guidance precision. The visual track linearization guidance method is suitable for on-line guidance of a superplatform.
Claims (1)
1. A multi-constraint on-line flight path planning middle section guiding method based on an air suction hypersonic speed platform is characterized by comprising the following specific steps:
(1) Online flight trajectory planning method based on predicted hit point
(1.1) air-breathing hypersonic platform flight strategy
After the air suction type hypersonic speed platform is launched, the air suction type hypersonic speed platform is subjected to two flight states of a boosting section and a scramjet power cruising section;
(a) Boosting section: the boosting section adopts a single-stage carrier rocket ascending section instruction generation rule; dividing the ascending process of the aircraft into a plurality of flight stages, setting fixed program angle instructions in each stage, determining the instructions by a limited number of parameters, performing off-line optimization according to task requirements to obtain the parameters and corresponding program instructions, and binding the parameters and the corresponding program instructions in a booster system, wherein the program instructions can be realized by using a gesture control system in the ascending stage flight process; to seek rapid generation of flight trajectory, the entire ascending segment is divided into 3 segments including vertical ascending segment 0.ltoreq.t.ltoreq.t 11 Negative angle of attack turning section T 11 ≤t≤T 12 Gravity turning section T 12 ≤t≤T 13 ；
The whole-course flight attack angle approximation formula of the suction hypersonic speed platform is as follows:
wherein: 1/v 0 Is the thrust-weight ratio; alpha m Is the maximum negative angle of attack magnitude, t m At the maximum moment of negative attack angle value, T 12 The negative attack angle turning section can be freely arranged; t (T) 13 The fuel exhaustion time of the solid rocket;
(b) Scramjet power cruise section: for an air suction type air suction hypersonic speed platform, the thrust provided by the scramjet engine is obtained through data interpolation:
wherein Q is dynamic pressure; ρ is the air density at the height of the aspirated hypersonic platform; k (k) r The opening of the valve is regulated for thrust; c is the sound velocity; ma is Mach number; t is engine thrust; v is the speed of the suction hypersonic speed platform;
(b.1) valve opening: because the cruising speed is higher than the speed at the end of the boosting section, the air suction type hypersonic speed platform is required to accelerate to a specified speed at first and then to be in a constant-speed flight state; the thrust is the maximum value provided by the scramjet engine in the accelerating flight process, namely the valve opening k r =1, i.e.:
T＝T max
in order to maintain a constant speed in a constant speed flight state, it is necessary to make
Namely:
wherein: θ, δ, m, r, x, y, z, α represent three components of the ballistics dip angle, ballistics deflection angle, mass, centroid distance and emission system of the aspiration hypersonic platform, respectively, and the attack angle; t, D the engine thrust and drag, respectively; g is the earth surface gravity acceleration;
the opening of the valve is as follows:
(b.2) angle of attack: dividing a flight track into a flat flight section and a climbing section based on the flight altitude; the attack angle instruction is obtained according to program instructions of different flight states;
(b.2.1) Flat fly section
The flat flight section requires the change rate of the altitude of the suction hypersonic platform to be unchanged, so that an aircraft is required
As a result of:
h＝r-R e
wherein: h. r is R e Respectively representing the flying height and the earth radius of the suction hypersonic platform;
deriving x, y, z and then summing the three components V of the transmission system speed x 、V y 、V z Deriving to obtain the second derivative change rate of the attack angle:
due to
Related to rise and resistance, which in turn are related to angle of attack, by setting an initial iteration value alpha of angle of attack 0 Small angle of attack Δα by iteration +.>
Obtaining an angle of attack by way of (a);
wherein: alpha k Is the attack angle value after iterating k times, and lambda is the optimal step factor of the damping Newton method, so that:
wherein: d, d k Is the search direction, order
Namely:
wherein:
wherein:
is the attack angle alpha k Acceleration of each axis of the emission coordinate system at the moment, +.>
Is the attack angle alpha k Acceleration of each axis of emission coordinate system at +Δα, setting cutoff condition |α k+1 -α k |<Epsilon, epsilon is small, then theObtaining the attack angle value of the flat flight segment when the cut-off condition is met;
(b.2.2) climbing section
The climbing section refers to a flight process of the suction hypersonic platform climbing to a designated height, and is divided into three parts, namely a fixed attack angle climbing section, a linear climbing section and a smooth transition section;
firstly, the suction hypersonic platform changes the inclination angle of a flight track by increasing lift force at a fixed attack angle climbing section, so that the suction hypersonic platform has climbing capacity;
when the inclination angle of the flight track reaches a preset value, the flight track ascends to a certain height in a linear climbing mode; in order to keep the intercepted flight path to climb straight, the inclination angle change rate of the flight path is required to be constant at 0 degrees/s, and the inclination angle change rate of the flight path is as follows:
setting an iteration initial value alpha of attack angle 0 By newton's iteration, namely:
by cyclic iteration of angle of attack alpha k To
Obtaining a flight attack angle alpha as an iteration termination condition, and obtaining a linear climbing flight track through integration;
finally, through a smooth transition section, a program instruction of the inclination angle of the flight track is obtained by taking the flight height as an independent variable and taking the inclination angle of the flight track as an excessive function of the dependent variable;
θ d (h)＝φ(h)*θ 1 +(1-φ(h))θ 0
wherein: h is a max To intercept the height, h min For a given smooth segment start height, h is the actual height, θ 0 To smooth the initial flight path inclination angle theta 1 To expect the inclination angle of the flight track, the inclination angle theta of the flight track is expected under the coordinate system of the flight track at the smooth tail end due to the design requirement of equal height and constant speed 1 =0°, the flight trajectory inclination command is reduced to:
θ d (h)＝(1-φ(h))θ 0
therefore, in order to convert the flight trajectory inclination angle instruction in the flight trajectory coordinate system into the flight trajectory inclination angle instruction in the emission system, it is necessary to obtain by conversion through the coordinate conversion matrix:
wherein: delta d As any one standard flight track group exists only in the longitudinal plane, so that v=delta d 0 deg. ≡0 °; after the flight track inclination angle theta' under the emission coordinate system is obtained, the variation rate of the flight track inclination angle is obtained based on the iteration step delta t by subtracting the flight track inclination angle theta of the current actual emission coordinate system from the flight track inclination angle theta of the current actual emission coordinate system:
the change rate of the inclination angle of the flight trajectory under the emission coordinate system is expressed as follows:
iteration of attack angles is completed through the change rate of the inclination angle of the flight track, and the attack angles are substituted into the dynamic model integration to obtain the flight track of the smooth section, and the specific method is the same as that of the flat flight section;
(1.2) design of on-line flight trajectory planning method
On-line design of a flight track based on the air-pushing coupling characteristic of the air-sucking platform, wherein control parameters are attack angle, tilting angle and engine throttle;
in the transverse plane, considering a yaw path, the yaw path turns based on a fixed roll angle, performs track design by eliminating a lead angle, and gives a roll angle instruction based on the direction of the lead angle:
wherein the method comprises the steps of
To allow maximum deviation of line of sight angle, gamma 0 For a set fixed roll angle magnitude;
in the longitudinal plane, a longitudinal flight strategy is designed based on the principle that fuel consumption is low and an equal dynamic pressure state can be maintained for a long time, and specifically comprises the following steps:
if the current flight height h is larger than the predicted hit point height, the characteristic of low flight track specific impulse is utilized, firstly, a downward flight track is adopted, the height of the predicted hit point is reduced, and then a cruise flight mode is adopted to fly against the predicted hit point;
if the current flight height h is smaller than the predicted hit point height, in order to avoid premature climbing, specific impact is reduced, fuel consumption is uneconomical, the aircraft flies at the current flight height, climbs to the interception height after the remaining flight time is smaller than a preset value, and returns to the cruising flight mode again;
(2) Nominal track tracking guidance method design
Obtaining a nominal track through the step (1), and then designing a proper guidance rule so that the actual flight track of the middle guide section well tracks the nominal track; the method comprises the following steps:
the trajectory dynamics model is dimensionally expanded by taking the speed, the altitude and the transverse position as state variables, and the state space is expressed as
Wherein x (t) is the state quantity of the middle guide section in the flight process; u (t) is a control variable;
assume that
In the method, in the process of the invention,
is a nominal track state quantity; />
Is a nominal trajectory control quantity; e is the difference between the actual flight trajectory and the nominal trajectory state quantity, +.>
The difference value between the control quantity and the nominal track control quantity in the actual flight process is obtained; then
Has the following components
Along with
Linearization, can obtain
In the method, in the process of the invention,
for the above system, a time-varying controller is used, which takes the form:
in the method, in the process of the invention,
wherein K (t) is a control parameter, a 3×6 matrix; substituting the above into a time-varying controller to obtain a closed-loop system matrix;
let the desired closed-loop matrix be
Wherein lambda is i (i=1, …, 6) is the desired feature root;
the closed-loop system matrix is equal to the closed-loop system matrix obtained above, and each control parameter in the expected closed-loop matrix can be obtained
In the method, in the process of the invention,
the control parameters are determined by the expected characteristic roots, so that proper expected characteristic roots need to be selected when the control parameters are required;
the three subspaces can be described as a second order system, described in terms of a characteristic equation, as follows:
the solution can be obtained:
in the method, in the process of the invention,
is natural frequency omega 1 (t),ω 2 (t),ω 3 (t) damping ratio, by demand performance to determine the desired system +.>
And omega 1 (t),ω 2 (t),ω 3 (t) thereby designing controller parameters that meet performance requirements;
for an underdamped second order linear system, the step response rise time and overshoot estimation formulas are as follows:
the rising time and overshoot of the system can be obtained by the above formula, and meanwhile, the damping ratio and the frequency of the expected system can be obtained, so that the characteristic root of the expected system matrix can be determined, and the controller parameters can be obtained.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310088829.0A CN116301028B (en) | 2023-02-09 | 2023-02-09 | Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310088829.0A CN116301028B (en) | 2023-02-09 | 2023-02-09 | Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN116301028A true CN116301028A (en) | 2023-06-23 |
| CN116301028B CN116301028B (en) | 2023-08-04 |
Family
ID=86823096
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310088829.0A Active CN116301028B (en) | 2023-02-09 | 2023-02-09 | Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116301028B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117234070A (en) * | 2023-11-13 | 2023-12-15 | 西安现代控制技术研究所 | BTT distribution method based on angle control instruction |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4465249A (en) * | 1981-04-01 | 1984-08-14 | Societe Nationale Industrielle Aerospatiale | Lateral acceleration control method for missile and corresponding weapon systems |
| CN106568355A (en) * | 2016-11-01 | 2017-04-19 | 湖北航天技术研究院总体设计所 | Missile-rotating satellite searching method for hypersonic velocity missile |
| CN110515392A (en) * | 2019-08-26 | 2019-11-29 | 哈尔滨工业大学 | A kind of hypersonic aircraft Trajectory Tracking Control method that performance oriented restores |
| CN114237299A (en) * | 2021-12-22 | 2022-03-25 | 北京航空航天大学 | Aircraft guidance method and device, electronic equipment and storage medium |
-
2023
- 2023-02-09 CN CN202310088829.0A patent/CN116301028B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4465249A (en) * | 1981-04-01 | 1984-08-14 | Societe Nationale Industrielle Aerospatiale | Lateral acceleration control method for missile and corresponding weapon systems |
| CN106568355A (en) * | 2016-11-01 | 2017-04-19 | 湖北航天技术研究院总体设计所 | Missile-rotating satellite searching method for hypersonic velocity missile |
| CN110515392A (en) * | 2019-08-26 | 2019-11-29 | 哈尔滨工业大学 | A kind of hypersonic aircraft Trajectory Tracking Control method that performance oriented restores |
| CN114237299A (en) * | 2021-12-22 | 2022-03-25 | 北京航空航天大学 | Aircraft guidance method and device, electronic equipment and storage medium |
Non-Patent Citations (1)
| Title |
|---|
| 王健权等: "基于高超声速飞机平台的拦截中制导方法研究", 空天技术, no. 6, pages 67 - 76 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117234070A (en) * | 2023-11-13 | 2023-12-15 | 西安现代控制技术研究所 | BTT distribution method based on angle control instruction |
| CN117234070B (en) * | 2023-11-13 | 2024-03-19 | 西安现代控制技术研究所 | BTT distribution method based on angle control instruction |
Also Published As
| Publication number | Publication date |
|---|---|
| CN116301028B (en) | 2023-08-04 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111306998B (en) | Parameter perturbation self-adaptive guided rocket projectile vertical attack guidance method | |
| CN103090728B (en) | Tail angle restraining guidance method based on sliding mode control | |
| CN111580547B (en) | Hypersonic aircraft formation control method | |
| CN111306989A (en) | Hypersonic velocity reentry guidance method based on steady glide trajectory analytic solution | |
| US11474540B2 (en) | Aircraft control method | |
| CN103983143B (en) | Downslide section method of guidance thrown in by the air-to-ground guided missile comprising rate process constraint and multiple terminals constraint | |
| CN109858106A (en) | Aircraft winglet stroke optimization method based on Gauss puppet spectrometry | |
| CN111336871B (en) | Vertical attack guidance method based on circuitous flight | |
| CN116301028B (en) | Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform | |
| CN110515392A (en) | A kind of hypersonic aircraft Trajectory Tracking Control method that performance oriented restores | |
| CN109240335A (en) | A kind of re-entry space vehicle approach method of guidance | |
| CN115079565B (en) | Variable-coefficient constraint guidance method and device with falling angle and aircraft | |
| CN114912202A (en) | Integrated coupling control method for propelling of wide-speed-range air-breathing power aircraft body | |
| CN113741509B (en) | Hypersonic gliding aircraft hold-down section energy management method | |
| CN111240204A (en) | Model reference sliding mode variable structure control-based flying bomb patrol control method | |
| CN114637312A (en) | Unmanned aerial vehicle energy-saving flight control method and system based on intelligent deformation decision | |
| CN113504723B (en) | Carrier rocket load shedding control method based on inverse reinforcement learning | |
| CN112660426B (en) | Rocket soft landing guidance method | |
| CN109101034B (en) | Flight control method for vertical/short-distance takeoff and landing aircraft | |
| CN116699994A (en) | Aircraft control method based on reinforcement learning | |
| CN114610057B (en) | Design method for maneuver burst prevention strategy of high Mach aircraft | |
| CN110209179A (en) | A kind of prompt high track algorithm of hypersonic aircraft | |
| CN115685764A (en) | Task self-adaptive anti-interference tracking control method and system for variable-span aircraft | |
| CN111506096B (en) | Tail-pushing type double-duct unmanned aerial vehicle fuzzy self-adaptive attitude control method | |
| CN111026153A (en) | Guiding method and guiding device for increasing flight distance of boosting gliding aircraft |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |