CN113969842B - High-dynamic-response control method for thrust of aero-engine - Google Patents

High-dynamic-response control method for thrust of aero-engine Download PDF

Info

Publication number
CN113969842B
CN113969842B CN202111235659.1A CN202111235659A CN113969842B CN 113969842 B CN113969842 B CN 113969842B CN 202111235659 A CN202111235659 A CN 202111235659A CN 113969842 B CN113969842 B CN 113969842B
Authority
CN
China
Prior art keywords
thrust
rotating speed
control
instruction
engine
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.)
Active
Application number
CN202111235659.1A
Other languages
Chinese (zh)
Other versions
CN113969842A (en
Inventor
李睿超
王向阳
王兵
张会强
谢峤峰
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Qinghang Aerospace Beijing Technology Co ltd
Tsinghua University
Original Assignee
Qinghang Aerospace Beijing Technology Co ltd
Tsinghua University
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Qinghang Aerospace Beijing Technology Co ltd, Tsinghua University filed Critical Qinghang Aerospace Beijing Technology Co ltd
Priority to CN202111235659.1A priority Critical patent/CN113969842B/en
Publication of CN113969842A publication Critical patent/CN113969842A/en
Application granted granted Critical
Publication of CN113969842B publication Critical patent/CN113969842B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
    • F02C9/48—Control of fuel supply conjointly with another control of the plant

Abstract

The invention provides a control method for high thrust dynamic response of an aeroengine, which has the core idea of avoiding slower dynamic responseIs changed. The control system based on the method comprises two steps of control plan (realized by a thrust instruction setting module 1, a rotating speed instruction setting module 2 and a control input reference value setting module 3) design and control algorithm (realized by a design controller 4) design: the module 1 generates a thrust instruction F according to the throttle lever angle and the air inlet condition cmd (ii) a Module 2 according to F cmd Generating a speed command N cmd (ii) a Module 3 according to F cmd And N cmd Generating fuel flow and geometrically variable component area/angle reference values; the controller 4 is a constant tracking feedback controller, and is designed to accelerate the response speed based on a pole allocation method and realize the non-static tracking of the instruction through a built-in integrator. The control method provided by the invention can enable the engine to have high dynamic response capability, and has potential application value in scenes such as stealth aircraft course control, carrier-based aircraft landing and the like.

Description

High-dynamic-response control method for thrust of aero-engine
Technical Field
The invention relates to a control method of an aero-engine, and belongs to the technical field of aero-engine control.
Background
The aircraft engine provides thrust for the aircraft, and the aircraft has cruising and maneuvering capabilities. However, the thrust cannot be directly measured, and the conventional engine control system can only indirectly control the thrust by controlling measurable parameters such as rotating speed and the like. With the improvement of the performance of the aircraft, the indirect control method cannot meet the fine requirement of the aircraft on the thrust, and the direct control problem of the thrust needs to be solved urgently.
From the point of view of the functioning of the propulsion system, the study of the direct control of thrust can be divided into two phases, with a slight overlap: the first phase is about 1978-1995, where the emphasis was on optimizing the performance indicators (acceleration, range, etc.) of aircraft by improving the overall performance (thrust, fuel consumption) of the engine, a typical work being engine performance optimization control arising in the nineties of the last century; the second phase was 1990 to date, where the focus was on using engines to control aircraft attitude, and where the dynamic behaviour of engines and aircraft was of interest compared to the first phase. Typical application scenes comprise unsteady maneuvering of a fighter plane, takeoff and landing of a short-vertical-takeoff and landing aircraft, carrier-based aircraft landing, course control of a high-stealth aircraft and the like. In these application scenarios, the thrust is not only controlled to be "steady" and "quasi", but also has a sufficiently "fast" response speed.
At present, there are three main technical schemes for improving the response speed of the engine at home and abroad: 1) Improving the bandwidth of the control system: the response speed is increased by increasing the bandwidth of the rotating speed control loop; 2) High-speed slow moving: the gas compressor is used for releasing gas to increase the rotating speed of the slow car, so that the acceleration time from the slow car to the maximum state is shortened; 3) Relaxing the restriction requirements on the engine: in an emergency situation, the limits on compressor surge margin and turbine front temperature are relaxed, allowing the fuel flow to change at a faster rate. However, these methods have the following limitations: 1) Mainly faces to civil aviation engines, and is not optimized aiming at the structure and the control characteristics of military aviation engines; 2) The starting point is still to improve the response speed of the rotor instead of directly controlling the thrust; 3) The method for high-speed slow vehicle and limitation relaxation is only suitable for a specific power state (such as slow vehicle) or a specific flight scene (such as takeoff), and cannot be used as a stock control method;
compared with the existing control method, the novel control method is provided aiming at the characteristic that the military aircraft engine is generally provided with the device with the adjustable throat area of the tail nozzle, the response speed of the thrust is effectively improved by the method of avoiding the fluctuation of the rotating speed, and the novel control method has good application value in aircraft attitude control scenes such as heading control of stealth aircraft, carrier landing of carrier-based aircraft and the like.
Disclosure of Invention
The invention aims to provide a thrust high-dynamic-response control method for an aeroengine. The method needs to complete control plan design and control algorithm design. The control plan specifies the state of the engine, which is realized by a thrust instruction setting module 1, a rotating speed instruction setting module 2 and a control input reference value setting module 3, and the control algorithm enables the engine to work in the specified state, which is realized by a controller 4.
Designing a control plan: the invention provides a novel control plan, which has the following specific functions: 1) The throttle lever angle is directly mapped to a thrust command: the method comprises the steps that the range of the required thrust of the aircraft is predetermined according to flight conditions, and the angle change range of the throttle lever is linearly mapped between the minimum value and the maximum value of the required thrust; 2) And determining a rotating speed instruction according to the thrust instruction, wherein the specific logic is as follows: when the thrust instruction changes, judging whether the current rotating speed meets the thrust requirement or not by looking up a table, if so, freezing the rotating speed instruction, enabling the engine to change the thrust along an equal rotating speed line, and after the thrust reaches a stable state, regenerating the rotating speed instruction to guide the engine to adjust the rotating speed to a proper value along the equal rotating speed line. When the rotating speed can not meet the thrust requirement, the rotating speed instruction is generated again, and at the moment, the thrust and the rotating speed of the engine are changed simultaneously. 3) Reference values of control inputs (including but not limited to fuel flow, throat area of a tail nozzle, guide vane angle, opening degree of an injection valve and the like) of the engine are generated according to a thrust instruction and a rotating speed instruction by means of table lookup.
Designing a control algorithm: the invention uses a constant tracking feedback controller to directly control the rotating speed and the thrust, and the controller 4 calculates a control signal according to a thrust instruction, a rotating speed instruction and a control input reference value to complete a control target.
Drawings
FIG. 1: the invention provides a schematic diagram of a thrust high dynamic response process.
FIG. 2: the invention provides a high-thrust dynamic response control system schematic block diagram.
FIG. 3: and a working logic diagram of the rotating speed setting module.
FIG. 4: the available thrust range and thrust radius are illustrated.
FIG. 5: the corresponding relation between the rotating speed instruction and the thrust instruction is shown in the figure.
FIG. 6: the invention is compared with the traditional scheme in the thrust response speed.
Detailed Description
The invention is further described with reference to the following figures and examples, which are included to provide a further understanding of the invention and are not to be construed as limiting the invention in any way.
Taking an engine as an example, an equal rotation speed line and an equal thrust line are taken, as shown in FIG. 1, and the control input in the figure is a fuel flow W F Throat area A of the rear nozzle 8 The rotation speed is a low-pressure rotor rotation speed line, and the principle of the control method is only illustrated here, and the type of the control input and the number of rotors are not particularly limited. Taking the initial thrust of 50% as an example, the acceleration process is kept unchanged compared with the traditional tail nozzle throat area (see A1->Path B1), the method provided by the invention enables the engine to be in the same initial thrust state along the path A2->B2->C2 path change, wherein A2->B2 is a constant rotation speed thrust increasing process, B2->C2 is a constant thrust and rotating speed increasing process. The functions of the respective sub-modules for realizing the control plan will be described below.
As shown in fig. 2, the control system based on this control method includes: the device comprises a thrust instruction setting module 1, a rotating speed instruction setting module 2, a control input reference value setting module 3 and a controller 4. The working logic or design method of each module is as follows:
1) The design method of the rotating speed instruction setting module 1 is as follows: the method comprises the steps that the range of the required thrust of the aircraft is determined in advance according to flight conditions, and the angle change range of the throttle lever is linearly mapped between the minimum value and the maximum value of the required thrust;
2) The operating logic of the rotational speed instruction setting module 2 is shown in fig. 3, and the implementation method of the decision logic or each flow of each step is as follows:
Figure BDA0003317592230000033
judging whether the rotating speed meets the thrust requirement: the method comprises the steps of changing each actuating mechanism (including but not limited to fuel flow, throat area of a tail nozzle, guide vane angle, opening degree of an injection valve and the like) of the engine in an actuating range, and recording thrust corresponding to each rotating speed under the condition that engine parameters are not over-limitThe range of variation, and referred to as the available thrust range, is shown in FIG. 4. If the thrust instruction is in the available thrust range corresponding to the current rotating speed, the rotating speed meets the thrust requirement, otherwise, the rotating speed does not meet the thrust requirement;
Figure BDA0003317592230000034
judging whether the thrust reaches a steady state: when the relative error e between the thrust command and the actual thrust is less than epsilon, the duration is more than T ε The thrust is considered to reach a steady state, where ε, T ε Is a preset constant.
Figure BDA0003317592230000035
The method for generating the rotating speed instruction comprises the following steps: the minimum available thrust force F at a certain speed of rotation is obtained according to FIG. 4 min And the maximum available thrust F max . Given thrust command F cmd Defining the thrust radius R at a certain speed as:
R＝min(F cmd –F min ,F max –F cmd )
the definition if and only if F min <F cmd <F max Is present. The rotating speed at which the radius of thrust is maximized is recorded as
Figure BDA0003317592230000031
Thrust command F cmd And with
Figure BDA0003317592230000032
See fig. 5 for an example of the correspondence relationship of (a). FIG. 5 also shows the thrust-to-rotation speed correspondence relationship obtained by minimizing the throat area of the nozzle tip or minimizing the fuel consumption rate in the conventional control scheme, in which the rotation speed is recorded as
Figure BDA0003317592230000041
Generally, the conventional control plan has a smaller fuel consumption rate, and therefore, a weight coefficient α defined between 0 and 1 generates a rotation speed command as follows:
Figure BDA0003317592230000042
the rotating speed command generated by the method can balance the control effect between the thrust response speed and the economic index. Similarly, the control effect may be balanced between thrust response speed and other indicators such as turbine inlet temperature (turbine life).
3) The control input reference value setting module 3 generates the control input reference value by the following method: tabulating the corresponding relation between the control input and the thrust and the rotating speed in the figure 1, and tabulating the table by the module according to the thrust instruction and the rotating speed instruction to obtain a control input reference value;
4) The controller 4 is designed as follows: for a given steady state operating condition (F) cmd ,N cmd ,W f,ref ,A 8,ref Check a), the engine linearization model is:
Figure BDA0003317592230000043
δy＝C·δx+D·δu
where δ represents the difference between the variable and its steady state value, the state quantity x is the rotor speed N, x = [ N ] when there are multiple rotors 1 ,N 2 ,...] T Control amount u = [ W ] f ,A 8 ,...] T Output y = [ F, N =] T When there are a plurality of rotors, y = [ F, N = 1 ,N 2 ,...] T 。
Defining errors
Figure BDA0003317592230000044
And in conjunction with the engine linearization equation, the following augmented linearization equation can be obtained:
Figure BDA0003317592230000045
ignoring actuator dynamics for the augmentation systemSystem for designing state feedback controller by using pole allocation method
Figure BDA0003317592230000046
And block-wise expand K and
Figure BDA0003317592230000047
the mathematical expression for the controller can be found as:
u＝K x (x-x ref )+K e ∫(y-y cmd )dt+u ref
in the design process, the response speed of the controlled object is increased by a method of enabling the pole to be far away from the virtual axis, and finally the pole position is determined according to the speed limit of the executing mechanism.
The controller is designed by using the method, compared with the traditional control scheme of only adjusting fuel, the poles of the two schemes are arranged in the same area, the control effect comparison that the thrust is increased from 50% to 60% is given in figure 6, and the thrust change curve at the upper right corner shows that the fuel oil control method has obvious response speed advantage.

Claims (1)

1. The high dynamic response control method of the aircraft engine thrust is characterized by comprising a control plan design and a control algorithm design; the control plan is to specify the state of the engine, and is realized by a thrust instruction setting module (1), a rotating speed instruction setting module (2) and a control input reference value setting module (3), and the control algorithm enables the engine to work in the specified state and is realized by a controller (4);
the control plan design specifically comprises:
1) The throttle lever angle is directly mapped to a thrust command: the thrust instruction setting module (1) determines the range of the required thrust of the aircraft according to the flight condition, and linearly maps the angle change range of the throttle lever between the minimum value and the maximum value of the required thrust;
2) And determining a rotating speed instruction according to the thrust instruction, wherein the specific logic is as follows:
judging whether the rotating speed meets the thrust requirement: changing each actuating mechanism of the engine within the actuating range, recording the thrust variation range corresponding to each rotating speed under the condition that the engine parameters do not exceed the limit, and calling the thrust variation range as an available thrust range; if the thrust instruction is in the available thrust range corresponding to the current rotating speed, the rotating speed meets the thrust requirement, otherwise, the rotating speed does not meet the thrust requirement;
judging whether the thrust reaches a steady state: when the relative error e between the thrust command and the actual thrust is less than epsilon for a time period greater than
Figure DEST_PATH_IMAGE002
The thrust is considered to reach a steady state, where ε,
Figure 860867DEST_PATH_IMAGE002
Is a preset constant;
the method for generating the rotating speed instruction comprises the following steps: obtaining the minimum available thrust F at a certain speed min And the maximum available thrust F max Given a thrust command F cmd Defining the thrust radius R at a certain rotational speed as: r = min (F) cmd –F min ,F max –F cmd ) The definition if and only if min <F cmd <F max When existing, the rotating speed which makes the thrust radius maximum is recorded as
Figure DEST_PATH_IMAGE004
The corresponding relation between the thrust and the rotating speed under the traditional control plan is obtained by keeping the throat area of the tail nozzle to be minimum or keeping the fuel consumption rate to be minimum, and the rotating speed under the control plan is recorded as
Figure DEST_PATH_IMAGE006
Defining a weight coefficient alpha between 0 and 1, and generating a rotating speed command according to the following formula:
Figure DEST_PATH_IMAGE008
；
3) The control input reference value of the engine is generated by looking up a table according to a thrust instruction and a rotating speed instruction: the control input reference value setting module (3) generates the control input reference value by the following method: tabulating the corresponding relation between the control input and the thrust and the rotating speed, and tabulating the table by the module according to the thrust instruction and the rotating speed instruction to obtain a control input reference value;
the control algorithm is specifically designed as follows: directly controlling the rotating speed and the thrust by using a constant tracking feedback controller, and calculating a control signal by using a controller (4) according to a thrust instruction, a rotating speed instruction and a control input reference value to finish a control target; the controller (4) is a constant tracking multiple input-multiple output feedback controller, and the design method is as follows: for a given steady state operating condition, the engine linearization model is:
Figure DEST_PATH_IMAGE010
where δ represents the difference between the variable and its steady state value, the state quantity x is the rotor speed N, and the control quantity u = [ W = f ,A 8 ] T Output y = [ F, N =] T Definition error of
Figure DEST_PATH_IMAGE012
And in conjunction with the engine linearization equation, the following augmented linearization equation can be obtained:
Figure DEST_PATH_IMAGE014
ignoring actuator dynamics, designing a state feedback controller using a pole placement method for the augmented linearization equation
Figure DEST_PATH_IMAGE016
And block-wise expand K and
Figure DEST_PATH_IMAGE018
the mathematical expression for the controller can be found as:
Figure DEST_PATH_IMAGE020
during design, the dynamic performance of the controller is accelerated by arranging the pole at a position far away from the virtual axis, the pole position is determined according to the speed limit of the actuating mechanism, the controller simultaneously controls the rotating speed and the thrust, and the static error-free tracking of rotating speed and thrust instructions is realized through a built-in integrator;
the method is based on the principle of avoiding the fluctuation of the rotating speed of the rotor to adjust the thrust of the aircraft engine, and the rotating speed instruction is determined by maximizing the thrust variation range under the constant rotating speed, so that the response speed and the thrust radius of the thrust are improved.
CN202111235659.1A 2021-10-22 2021-10-22 High-dynamic-response control method for thrust of aero-engine Active CN113969842B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202111235659.1A CN113969842B (en) 2021-10-22 2021-10-22 High-dynamic-response control method for thrust of aero-engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202111235659.1A CN113969842B (en) 2021-10-22 2021-10-22 High-dynamic-response control method for thrust of aero-engine

Publications (2)

Publication Number Publication Date
CN113969842A CN113969842A (en) 2022-01-25
CN113969842B true CN113969842B (en) 2022-11-18

Family

ID=79588033

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202111235659.1A Active CN113969842B (en) 2021-10-22 2021-10-22 High-dynamic-response control method for thrust of aero-engine

Country Status (1)

Country Link
CN (1) CN113969842B (en)

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4425614A (en) * 1969-12-09 1984-01-10 Adaptronics, Inc. Self-organizing control system
US6459963B1 (en) * 2000-07-31 2002-10-01 General Electric Company Methods and apparatus for trimming engine control systems
CN112855346A (en) * 2019-11-27 2021-05-28 中国航发商用航空发动机有限责任公司 Thrust control method and system and aircraft engine

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6171055B1 (en) * 1998-04-03 2001-01-09 Aurora Flight Sciences Corporation Single lever power controller for manned and unmanned aircraft
US6269627B1 (en) * 1998-12-16 2001-08-07 United Technologies Corporation Rapid thrust response control logic for shaft-driven lift fan STOVL engine
US7010904B2 (en) * 2003-12-05 2006-03-14 United Technologies Corporation Method for control scheduling to achieve linear thrust response
GB201321472D0 (en) * 2013-12-05 2014-01-22 Rolls Royce Plc Control of a gas turbine engine
US20200247552A1 (en) * 2019-02-06 2020-08-06 Pratt & Whitney Canada Corp. System and method for controlling propeller-driven aircraft
CN110513199B (en) * 2019-07-23 2021-06-22 南京航空航天大学 Neural network control method for relieving thrust recession of aircraft engine
CN111425304B (en) * 2020-04-23 2021-01-12 南京航空航天大学 Aero-engine direct thrust control method based on composite model predictive control
CN111852662A (en) * 2020-06-15 2020-10-30 西北工业大学 Fault-tolerant two-degree-of-freedom H-infinity controller for maximum thrust state of aircraft engine
CN112947064A (en) * 2021-01-21 2021-06-11 西北工业大学 Aero-engine maximum thrust control optimization method considering gas circuit component faults

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4425614A (en) * 1969-12-09 1984-01-10 Adaptronics, Inc. Self-organizing control system
US6459963B1 (en) * 2000-07-31 2002-10-01 General Electric Company Methods and apparatus for trimming engine control systems
CN112855346A (en) * 2019-11-27 2021-05-28 中国航发商用航空发动机有限责任公司 Thrust control method and system and aircraft engine

Also Published As

Publication number Publication date
CN113969842A (en) 2022-01-25

Similar Documents

Publication Publication Date Title
EP2154588B1 (en) Control system
EP3357810A1 (en) Power demand anticipation systems for rotorcraft
US20080255713A1 (en) Dynamic Adjustment of Wing Surfaces for Variable Camber
CN110083869B (en) Calculation method for evaluating influence of mode conversion on stability margin of turbojet/turbofan variable cycle engine
US20070110577A1 (en) Control logic for a propeller system
US4232515A (en) Supersonic cruise airplane and engine
US9322341B2 (en) System and method for engine transient power response
EP3748149A1 (en) Engine and thrust control of aircraft in no dwell zone
US4947643A (en) Active geometry control system for gas turbine engines
CN111731490B (en) Helicopter comprehensive control method and device with variable rotor rotation speed
US6269627B1 (en) Rapid thrust response control logic for shaft-driven lift fan STOVL engine
CN103003529A (en) Method and system for controlling the clearance at the blade tips of a turbine rotor
US20200271074A1 (en) Anti-unstart for combined cycle high mach vehicles
CN110374750A (en) A kind of aero-engine meets with the control method and device of acceleration fuel feeding
US9822731B2 (en) Control scheme using variable area turbine and exhaust nozzle to reduce drag
JP2923311B2 (en) Output control device for aircraft engine
CN113969842B (en) High-dynamic-response control method for thrust of aero-engine
US4523603A (en) Air intake control for an adjustable air inlet, particularly two-dimensional oblique shock diffuser for gas turbine jet engines for the propulsion of high performance aircraft
CN116301028B (en) Multi-constraint on-line flight trajectory planning middle section guiding method based on air suction hypersonic speed platform
CN112231835A (en) Thrust performance and deflection efficiency integrated vectoring nozzle outlet area optimization method
CN104500267B (en) For regulating the fan-shaped throttling arrangement of turbine electricity generation system power
KR100569766B1 (en) Method of operating a gas turbine engine
Bates et al. Design and piloted simulation of a robust integrated flight and propulsion controller
CN112512923A (en) Aircraft propulsion parameter adjustment method and related computer program
Bates et al. Integrated flight and propulsion control system design using H/sup/spl infin//loop-shaping techniques

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