CN111597633A - Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon - Google Patents

Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon Download PDF

Info

Publication number
CN111597633A
CN111597633A CN202010386074.9A CN202010386074A CN111597633A CN 111597633 A CN111597633 A CN 111597633A CN 202010386074 A CN202010386074 A CN 202010386074A CN 111597633 A CN111597633 A CN 111597633A
Authority
CN
China
Prior art keywords
pylon
design
vibration
loop
omega
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.)
Pending
Application number
CN202010386074.9A
Other languages
Chinese (zh)
Inventor
王继强
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.)
Nanjing University of Aeronautics and Astronautics
Original Assignee
Nanjing University of Aeronautics and Astronautics
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 Nanjing University of Aeronautics and Astronautics filed Critical Nanjing University of Aeronautics and Astronautics
Priority to CN202010386074.9A priority Critical patent/CN111597633A/en
Publication of CN111597633A publication Critical patent/CN111597633A/en
Pending legal-status Critical Current

Links

Images

Classifications

    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/17—Mechanical parametric or variational design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces
    • 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)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Theoretical Computer Science (AREA)
  • General Physics & Mathematics (AREA)
  • Evolutionary Computation (AREA)
  • Computer Hardware Design (AREA)
  • General Engineering & Computer Science (AREA)
  • Pure & Applied Mathematics (AREA)
  • Mathematical Optimization (AREA)
  • Mathematical Analysis (AREA)
  • Computational Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Vibration Prevention Devices (AREA)

Abstract

The invention discloses a rigidity feedback design method for coupling vibration reduction of an aero-engine and a pylon, and belongs to the field of vibration reduction of aero-engines. The method comprises the following steps: the method comprises the steps of establishing a coupling system model of the aircraft engine and a pylon; carrying out the open loop performance analysis of the coupling system; carrying out rigidity feedback design to obtain the optimal control parameters of the coupling system; simulations were performed to confirm that the optimal control met the performance requirements. The invention can obtain the optimal performance of the coupling system of the aircraft engine and the pylon.

Description

Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon
Technical Field
The invention relates to a rigidity feedback design method for coupling vibration reduction of an aero-engine and a pylon, belonging to the field of vibration reduction of aero-engines.
Background
The vibration of the aircraft engine is transmitted to the aircraft through the pylon and causes the aircraft body to flutter, which not only affects comfort, but also may cause structural fatigue and even serious accidents. However, the vibration of aircraft engines is extremely difficult to control, one of the reasons being the variety of sources that cause the vibration, for example, assembly errors, aerodynamic disturbances, combustion chamber excitation, and the like. The most significant of these is the aerodynamic imbalance of the high pressure rotor system (note that aerodynamic imbalance involves vibrations caused by aerodynamic disturbances, rotor misalignment, and bearing damage). The vibration caused by the pneumatic unbalance is transmitted to the outer duct casing through the core structure and then transmitted to the airplane through the hanging frame. Thus, there are generally two types of methods available to control vibration, one is to exert control from within the core machine to mitigate the effects of vibration from a source; and secondly, applying control on the transmission path to enable the vibration to be attenuated at the hanging point of the hanging frame.
Typically, aircraft engines and pylons are damped passively, i.e. by squeezing an oil film damper to absorb vibrational energy and dissipate the vibrational energy in the form of heat. Therefore, the passive mode vibration damping is to optimize the design of the squeeze film vibration damper or the corresponding support structure. For example, patent CN109720583A by airbus operating simplification companies discloses a main structure of an aircraft power plant support pylon; patent CN107108039B by lode corporation discloses an engine mounting system with failsafe attachment points that includes a front mounting bracket loaded with a coat hanger type latch and a retaining double wrench style shim.
However, such passive damping is more limited and required by the weight, size, space, etc. of the system, and to overcome these limitations, active damping control may be employed. However, no public report is found in the field of coupling vibration reduction of the aero-engine and the pylon at present.
Disclosure of Invention
The invention provides a rigidity feedback design method for coupling vibration reduction of an aero-engine and a pylon, aiming at the coupling system of the aero-engine and the pylon.
The invention adopts the following technical scheme for solving the technical problems:
a rigidity feedback design method for coupling vibration reduction of an aircraft engine and a pylon comprises the following steps:
(1) establishing a coupling system model of the aircraft engine and the pylon;
(2) carrying out the open loop performance analysis of the coupling system;
(3) carrying out rigidity feedback design to obtain the optimal control parameters of the coupling system;
(4) the optimal control is confirmed by simulation to meet the performance requirement; and (3) otherwise, returning to the steps (2) and (3), and determining whether the performance index reaches the expected value according to the performance limit of the coupling system.
The model of the coupling system of the aircraft engine and the pylon in the step (1) is as follows:
X1(jω)＝g11(jω)U(jω)+g12(jω)D(jω)
X2(jω)＝g21(jω)U(jω)+g22(jω)D(jω)；
U(jω)＝k2X2(jω)
wherein, X1(jω)、X2(j omega) is a frequency response function of two hanging points of the hanging rack and the aircraft engine; d (j omega) is the pneumatic unbalanced vibration of the high-pressure rotor system, and U (j omega) is the active vibration control force; g11(j ω) is X1Frequency transfer function of (j ω) with respect to U (j ω), g12(j ω) is X1Frequency transfer function of (j ω) with respect to D (j ω), g21(j ω) is X2Frequency transfer function of (j ω) with respect to U (j ω), g22(j ω) is X2(j ω) frequency transfer function with respect to D (j ω); k is a radical of2I.e. the stiffness coefficient to be designed.
The specific process of the step (2) is as follows:
firstly, determining D (j omega) according to the frequency and amplitude of the pneumatic unbalanced vibration of the high-pressure rotor system;
second, g is obtained by system identification or linearization of the component-level model11(jω)、g12(jω)、g21(jω)、g22A specific expression of (j ω);
again, when the system is open-loop, i.e. no control input U (j ω) is 0 or k 20, amount of vibration | X at two hanging points1(j ω) | and | X2(j ω) | is:
Figure BDA0002483877180000032
the above equation defines the open-loop performance of the coupled system, and accordingly, the closed-loop design must be such that the open-loop performance is improved, i.e., | X1(j ω) | and | X2(j omega) | BiMust be reduced at the same time.
The system identification includes, for example, a frequency sweep method.
The specific process of the step (3) is as follows:
first, the following expression is calculated:
Figure BDA0002483877180000031
secondly, the method comprises the following steps: calculate | G (j ω) -1| if | G (j ω) -1| is > 1, then | X1The closed loop of (j omega) is optimally performed as
Figure BDA0002483877180000041
Wherein:2is desired | X2Closed loop design criteria of (j ω) |, e.g. requirement | X2(j omega) | closed loop is reduced by 3dB, then2＝0.707；
And then, | X2The closed loop of (j omega) is optimally performed as
(|1-G(jω)|-1|G(jω)|)；
Wherein:1is desired | X1Closed loop design criteria of (j ω) |, e.g. requirement | X1The (j omega) | closed loop is reduced by 6dB, then1＝0.5；
In the above calculation, if | G (j ω) -1| ≦ 1 occurs, then placing the optimal design at the (-G (j ω)) point yields | X2(j ω) | closed-loop vibration is attenuated to zero; and placed at the (-1, 0) point to obtain | X1(j ω) | damping to zero design result of closed-loop vibration;
selecting the value of the optimal design point according to the specific design requirement, setting the value as α (j omega), and obtaining the optimal design k under the corresponding optimal performance according to the following formula2：
Figure BDA0002483877180000042
The invention has the following beneficial effects:
aiming at the coupling system of the aero-engine and the pylon, the coupling system of the aero-engine and the pylon can achieve the limit performance of the coupling system of the aero-engine and the pylon through the active feedback design of rigidity.
Drawings
FIG. 1 is a stiffness feedback design flow chart for damping vibration of an aircraft engine and pylon coupling system.
FIG. 2(a) shows the vibration | X of the coupled system at the hanging point 11Schematic of (j ω) |; FIG. 2(b) shows the vibration | X of the coupled system at the hanging point 22(j ω) | schematic.
Detailed Description
The present invention is further illustrated by the following description in conjunction with the accompanying drawings and the specific embodiments, it is to be understood that these examples are given solely for the purpose of illustration and are not intended as a definition of the limits of the invention, since various equivalent modifications will occur to those skilled in the art upon reading the present invention and fall within the limits of the appended claims.
A rigidity feedback design method for coupling vibration reduction of an aircraft engine and a pylon is disclosed, and comprises the following steps according to the flow shown in figure 1:
step 1: establishing model of aircraft engine and pylon coupling system
A generic model of an aircraft engine and pylon coupling system can be expressed as:
Figure BDA0002483877180000051
wherein, X1(jω)、X2(j omega) is a frequency response function of two hanging points of the hanging rack and the aircraft engine; d (j omega) is the pneumatic unbalanced vibration of the high-pressure rotor system, and U (j omega) is the active vibration control force; g11(jω)、g12(jω)、g21(jω)、g22(j ω) are the respective frequency transfer functions; k is a radical of2I.e. the stiffness coefficient to be designed. That is, it is necessary to find the optimum design k2So that the amount of vibration | X at two hanging points1(j ω) | and | X2(j ω) | is reduced simultaneously (| · | represents taking the complex variable X1(j ω) and X2Absolute value of (j ω).
Note that: the above model is a generic model and corresponding active designs are all included in the patent claims.
Step 2: developing the open loop performance analysis of the coupling system model;
first, D (j ω) is determined based on the frequency and amplitude of the aerodynamic imbalance vibration of the high pressure rotor system, in the following example, a typical frequency of 1000Hz, i.e., ω06280 rpm, amplitude of 1mm as an example;
second, g is obtained by system identification (e.g., frequency sweep method) or linearization of the component-level model11(jω)、g12(jω)、g21(jω)、g22In the following embodiments, a frequency transfer function of a certain type is as follows:
Figure BDA0002483877180000061
wherein: Δ ═ 1.5- ω2+1.5jω](0.5-ω2+0.5jω)+(0.5+0.5jω)2. In this model, the frequency is normalized to a dimensionless unit ω of 1, i.e., the unit frequency corresponds to ω06280 rpm.
Again, when the system is open-loop, i.e. no control input U (j ω) is 0 or k 20, amount of vibration | X at two hanging points1(j ω) | and | X2(j ω) | is:
Figure BDA0002483877180000062
the value of (3) can be calculated by taking ω 1 into formula (2):
Figure BDA0002483877180000063
that is, (4) defines the open loop performance of the coupled system, and accordingly, the closed loop design must be such that the open loop performance is improved, i.e., | X1(j ω) | and | X2(j ω) | must be reduced simultaneously. It is noted that almost all current control design methods only consider | X |1(j ω) | (complementary sensitivity function) decrease or | X2The (j ω) | (sensitivity function) decreases, and the problem of how to design both decreases is not considered.
And step 3: calculating to obtain optimal control parameters of the rigidity feedback design;
first, the following expression is calculated:
Figure BDA0002483877180000071
wherein: g (j omega) is a complex variable;
according to (2), the following can be calculated:
Figure BDA0002483877180000072
secondly, the method comprises the following steps: calculate | G (j ω) -1| if | G (j ω) -1| is > 1, then | X1The closed loop of (j omega) is optimally performed as
Figure BDA0002483877180000073
Wherein:2is desired | X2Closed loop design criteria of (j ω) |, e.g. requirement | X2(j omega) | closed loop is reduced by 3dB, then2＝0.707。
And then, | X2The closed loop of (j omega) is optimally performed as
(|1-G(jω)|-1|G(jω)|) (8)
Wherein:1is desired | X1Closed loop design criteria of (j ω) |, e.g. requirement | X1The (j omega) | closed loop is reduced by 6dB, then1＝0.5。
In the above calculation, if | G (j ω) -1| ≦ 1 occurs, then placing the optimal design at the (-G (j ω)) point may result in | X2(j ω) | closed-loop vibration is attenuated to zero; and placing at (-1, 0) point to obtain | X1Selecting the value of an optimal design point according to specific design requirements, setting the value as α (j omega), and obtaining the optimal design k under the corresponding optimal performance according to the following formula2：
Figure BDA0002483877180000074
Wherein: alpha (j omega) is the value of the optimal design point selected according to the specific design requirement;
in the above embodiment, since | G (j ω) -1| ═ 2 > 1, | X can be obtained1The closed loop optimal performance of (j omega) is (7); and | X2The closed loop optimum performance of (j ω) | is (8). Now assume that the design target is the amount of vibration | X at two hang points1(j ω) | and | X2(j ω) | are all reduced by 3dB at the same time, i.e.1＝2When the value is 0.707, | X can be calculated from (7) and (8)1(j ω) | and | X2The closed-loop limit performance of (j ω) | is:
Figure BDA0002483877180000081
|1-G(jω)|-1|G(jω)|＝0.419
that is, | X1The closed loop limit performance of (j ω) | is 0.579, i.e., -4.75 dB; and | X2The closed loop limit performance of (j ω) | is 0.419, i.e., -7.56 dB. Now suppose that the design requirement is | X2(j omega) | closed loop is reduced by 3dB, then20.707; and need to convert | X1(j ω) | closed loop pushes to its performance limit, i.e. | X1If the closed loop of (j ω) | is reduced by 4.75dB, α (j ω) — 1+0.75j needs to be taken, which can be obtained from (9):
Figure BDA0002483877180000082
and 4, step 4: numerical simulation to confirm that the optimal design meets performance requirements;
if the design requirements are met, implementing the design; and otherwise, returning to the steps (2) and (3) to confirm the optimal performance of the suspension system under the condition of performance limit. If the requirements are still not met, the performance index requirement is over high, and the index requirement needs to be reduced. For the above embodiment, if the index requirement is | X1(j ω) | and | X2(j ω) | is reduced by 3dB or less, the design can be confirmed; and if the index requirement is | X2(j ω) | decreases by 3dB and | X1(j ω) | decreases by more than 4.75 dB; or | X1(j ω) | decreases by 3dB and | X2(j ω) | decreases by more than 7.56 dB. In view of | X1The closed-loop limit performance of (j ω) | is-4.75 dB and | X2The closed loop limit performance of (j ω) | is-7.56 dB (see equation (10)), which indicates that the performance index requirement must be lowered due to the excessively high index requirement. In fact, | X will appear if the adherence index is not changed1(j ω) | or | X2(j ω) | vibration enhancement, i.e., a degraded performance condition, will enhance the transmission of vibration from the aircraft engine to the aircraft through the pylon, thereby causing aircraft flutter and compromising comfort. Thus, suitable index requirements are now identified, defined as | X as design requirements2(j ω) | closed loop reduction 3dB ((j ω) |)20.707), and | X1(j ω) | closed loop pushes to its performance limit, i.e. | X1The (j ω) | closed loop is reduced by 4.75dB, and the real-time simulation results are shown in fig. 2(a) and fig. 2 (b). It can be seen that the designed closed-loop control does enable | X1(j ω) | and | X2The (j ω) | closed-loop vibration amount reaches the target of simultaneous reduction.
In a word, aiming at the coupling system of the aero-engine and the pylon, the coupling system of the aero-engine and the pylon achieves the limit performance through the active feedback design of rigidity; the design method can judge whether the performance index requirement is feasible or not, and has important guidance for actual engineering.

Claims (5)

1. A rigidity feedback design method for coupling vibration attenuation of an aircraft engine and a pylon is characterized by comprising the following steps:
(1) establishing a coupling system model of the aircraft engine and the pylon;
(2) carrying out the open loop performance analysis of the coupling system;
(3) carrying out rigidity feedback design to obtain the optimal control parameters of the coupling system;
(4) the optimal control is confirmed by simulation to meet the performance requirement; and (3) otherwise, returning to the steps (2) and (3), and determining whether the performance index reaches the expected value according to the performance limit of the coupling system.
2. The stiffness feedback design method for coupling vibration damping of the aircraft engine and the pylon according to claim 1, wherein the model of the coupling system of the aircraft engine and the pylon in the step (1) is as follows:
X1(jω)＝g11(jω)U(jω)+g12(jω)D(jω)
X2(jω)＝g21(jω)U(jω)+g22(jω)D(jω)；
U(jω)＝k2X2(jω)
wherein, X1(jω)、X2(j omega) is a frequency response function of two hanging points of the hanging rack and the aircraft engine; d (j omega) is the pneumatic unbalanced vibration of the high-pressure rotor system, and U (j omega) is the active vibration control force; g11(j ω) is X1Frequency transfer function of (j ω) with respect to U (j ω), g12(j ω) is X1Frequency transfer function of (j ω) with respect to D (j ω), g21(j ω) is X2Frequency transfer function of (j ω) with respect to U (j ω), g22(j ω) is X2(j ω) frequency transfer function with respect to D (j ω); k is a radical of2I.e. the stiffness coefficient to be designed.
3. The stiffness feedback design method for coupling vibration damping of the aero-engine and the pylon as claimed in claim 2, wherein the specific process of the step (2) is as follows:
firstly, determining D (j omega) according to the frequency and amplitude of the pneumatic unbalanced vibration of the high-pressure rotor system;
second, g is obtained by system identification or linearization of the component-level model11(jω)、g12(jω)、g21(jω)、g22A specific expression of (j ω);
again, when the system is open-loop, i.e. no control input U (j ω) is 0 or k20, amount of vibration | X at two hanging points1(j ω) | and | X2(j ω) | is:
Figure FDA0002483877170000023
the above equation defines the open-loop performance of the coupled system, and accordingly, the closed-loop design must be such that the open-loop performance is improved, i.e., | X1(j ω) | and | X2(j ω) | must be reduced simultaneously.
4. The stiffness feedback design method for coupled damping of an aircraft engine and pylon of claim 3 wherein the system identification comprises a frequency sweep method.
5. The stiffness feedback design method for coupling vibration damping of the aero-engine and the pylon according to claim 3, wherein the specific process of the step (3) is as follows:
first, the following expression is calculated:
Figure FDA0002483877170000021
secondly, the method comprises the following steps: calculate | G (j ω) -1| if | G (j ω) -1| is > 1, then | X1The closed loop of (j omega) is optimally performed as
Figure FDA0002483877170000022
Wherein:2is desired | X2Closed loop design criteria of (j ω) |, e.g. requirement | X2(j omega) | closed loop is reduced by 3dB, then2＝0.707；
And then, | X2The closed loop of (j omega) is optimally performed as
(|1-G(jω)|-1|G(jω)|)；
Wherein:1is desired | X1Closed loop design criteria of (j ω) |, e.g. requirement | X1The (j omega) | closed loop is reduced by 6dB, then1＝0.5；
In the above calculation, if | G (j ω) -1| ≦ 1 occurs, then placing the optimal design at the (-G (j ω)) point yields | X2(j ω) | closed-loop vibration is attenuated to zero; and placed at the (-1, 0) point to obtain | X1(j ω) | damping to zero design result of closed-loop vibration;
selecting an optimal design point according to specific design requirementsIs set to α (j ω), the optimum design k for the corresponding optimum performance is found according to the following equation2：
Figure FDA0002483877170000031
CN202010386074.9A 2020-05-09 2020-05-09 Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon Pending CN111597633A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202010386074.9A CN111597633A (en) 2020-05-09 2020-05-09 Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202010386074.9A CN111597633A (en) 2020-05-09 2020-05-09 Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon

Publications (1)

Publication Number Publication Date
CN111597633A true CN111597633A (en) 2020-08-28

Family

ID=72189445

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202010386074.9A Pending CN111597633A (en) 2020-05-09 2020-05-09 Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon

Country Status (1)

Country Link
CN (1) CN111597633A (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112881025A (en) * 2021-01-12 2021-06-01 南京航空航天大学 Method for vibration control and energy collection of aircraft engine

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20170364131A1 (en) * 2016-06-20 2017-12-21 Telefonaktiebolaget Lm Ericsson (Publ) System and method to enhance a feedback loop of a power converter
CN110502787A (en) * 2019-07-16 2019-11-26 南京航空航天大学 Quasi- zero stiffness damper optimum design method
CN111123705A (en) * 2019-12-18 2020-05-08 南京航空航天大学 Design method for active vibration control of propeller and transmission shaft system

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20170364131A1 (en) * 2016-06-20 2017-12-21 Telefonaktiebolaget Lm Ericsson (Publ) System and method to enhance a feedback loop of a power converter
CN110502787A (en) * 2019-07-16 2019-11-26 南京航空航天大学 Quasi- zero stiffness damper optimum design method
CN111123705A (en) * 2019-12-18 2020-05-08 南京航空航天大学 Design method for active vibration control of propeller and transmission shaft system

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112881025A (en) * 2021-01-12 2021-06-01 南京航空航天大学 Method for vibration control and energy collection of aircraft engine
CN112881025B (en) * 2021-01-12 2022-06-10 南京航空航天大学 Method for vibration control and energy collection of aircraft engine

Similar Documents

Publication Publication Date Title
Peixin et al. Vibration analysis and control technologies of hydraulic pipeline system in aircraft: A review
US5199856A (en) Passive structural and aerodynamic control of compressor surge
Gysling et al. Dynamic control of centrifugal compressor surge using tailored structures
Brennan et al. Vibration control
Wu et al. Design of dry friction and piezoelectric hybrid ring dampers for integrally bladed disks based on complex nonlinear modes
CN107451355B (en) Shock absorber design method
WO2020248184A1 (en) Aircraft engine actuator fault-orientated disturbance-free switch-in fault-tolerant control method
Chahine et al. A comparative study of coupled and decoupled fan flutter prediction methods under variation of mass ratio and blade stiffness
CN111597633A (en) Rigidity feedback design method for coupling vibration reduction of aero-engine and pylon
CN111881530B (en) Vibration reduction optimization design method for aeroengine
CN110427677B (en) Rapid design method for elbow structure in external pipeline system of aircraft engine
CN110594344B (en) Zero-damping vibration absorber optimization design method
Guo et al. Dynamic characteristics of a shrouded blade with impact and friction
Carstens et al. Numerical investigation of nonlinear fluid-structure interaction in vibrating compressor blades
Depriest Aircraft engine attachment and vibration control
CN110175392B (en) Aero-engine vibration transmission path analysis method based on OTPA method and physical modeling
CN112881025B (en) Method for vibration control and energy collection of aircraft engine
Zhao et al. Investigation of actuator placement approaches for active vibration control in the aircraft engine
Tavasoli et al. Adaptive robust boundary control of coupled bending‐torsional vibration of beams with only one axis of symmetry
Liu et al. Dynamic reliability of aircraft hydraulic pipelines under random pressure pulsation and vibration
CN112580157A (en) Vibration reduction design method for internal and external casings of aero-engine under extremely low frequency condition
Heidari et al. H∞ and H 2 optimization procedures for optimal design of support parameters of a flexible rotor
Ivanov et al. Dynamic loads acting on engine frame elements after fan blade out event study
Wu et al. On the performance of wavy dry friction and piezoelectric hybrid flexible dampers
Swanson et al. Design and effectiveness evaluation of an active vibration isolation system for a commercial jet 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