CN116374177A - Airplane deicing system based on SMA intelligent material - Google Patents
Airplane deicing system based on SMA intelligent material Download PDFInfo
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- CN116374177A CN116374177A CN202310371323.0A CN202310371323A CN116374177A CN 116374177 A CN116374177 A CN 116374177A CN 202310371323 A CN202310371323 A CN 202310371323A CN 116374177 A CN116374177 A CN 116374177A
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- 238000009434 installation Methods 0.000 claims description 12
- 238000001208 nuclear magnetic resonance pulse sequence Methods 0.000 claims description 9
- 238000010438 heat treatment Methods 0.000 claims description 8
- 229910000734 martensite Inorganic materials 0.000 claims description 8
- 230000008859 change Effects 0.000 claims description 7
- 238000000034 method Methods 0.000 claims description 6
- 238000006073 displacement reaction Methods 0.000 claims description 5
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- 229910001566 austenite Inorganic materials 0.000 claims description 4
- 238000011156 evaluation Methods 0.000 claims description 4
- 239000002520 smart material Substances 0.000 claims description 4
- 229910045601 alloy Inorganic materials 0.000 claims description 3
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
- B64D15/20—Means for detecting icing or initiating de-icing
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
- B64D15/12—De-icing or preventing icing on exterior surfaces of aircraft by electric heating
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
- B64D15/12—De-icing or preventing icing on exterior surfaces of aircraft by electric heating
- B64D15/14—De-icing or preventing icing on exterior surfaces of aircraft by electric heating controlled cyclically along length of surface
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Abstract
The invention discloses an airplane deicing system based on an SMA intelligent material, which relates to the technical field of airplane deicing and comprises the following components: the driving control system is used for designing a scheme of an SMA self-sensing closed-loop control system and realizing reliable, efficient and simple SMA driving control; the skin structure is designed into a new embedded SMA deicing skin structure, and the theoretical division of the structure is given; the self-adaptive deicing rapid reconstruction algorithm is designed to achieve the aims of improving the performance of the reconstruction algorithm and improving the control accuracy, and the SMA self-perception type closed-loop control system comprises resistance-temperature-deformation modeling, multi-pulse driving-feedback scheme design and closed-loop control algorithm design. According to the invention, on the basis of setting the SMA self-sensing closed-loop control technology, an SMA deicing technology study based on an embedded SMA deicing skin structure and an SMA driving control technology is carried out, so that the efficiency of an aircraft deicing system is improved, the power consumption is reduced, and meanwhile, the size and the weight of the deicing system are reduced, so that the intelligent aircraft deicing system is suitable for wide popularization and application.
Description
Technical Field
The invention relates to the technical field of aircraft deicing, in particular to an aircraft deicing system based on an SMA intelligent material.
Background
Aircraft icing is an important factor affecting the safe flight of an aircraft. If the front edge of the wing of the airplane is frozen, the weight of the airplane is increased, the lift force of the airplane is reduced, the resistance of the airplane is increased, the aerodynamic performance of the airplane is deteriorated, and the safety and the maneuvering performance of the airplane are finally affected. Icing of an aircraft tail can cause an aircraft to suddenly stall, and in any rapidly icing weather, the icing speed on the tail surface is faster than that of the side wings due to the fact that the surface area of the aircraft tail is smaller than that of the side wings. The accident of air accident caused by aircraft icing is huge, which not only causes the aircraft to destroy the people and death, but also wastes a great deal of manpower and financial resources. In order to solve the problem of aircraft icing, various countries have been striving to study the technology of preventing and removing ice. At present, two main measures are adopted for the problem of aircraft icing, namely an anti-icing technology for preventing ice from forming on the surface of the aircraft. And secondly, deicing technology, namely removing the existing ice layer on the surface of the aircraft.
However, in the prior art, SMA (shape memory alloy) is directly laid on the surface of the skin of the aircraft or laminated inside the skin, and this method tends to affect the aerodynamic shape of the aircraft, and large area SMA material needs to be laid, which increases the deicing power consumption and costs a lot in use.
Therefore, the existing aircraft deicing system based on SMA intelligent materials cannot meet the requirements in practical use, so there is an urgent need in the market for improved technology to solve the above problems.
Disclosure of Invention
The invention aims to provide an airplane deicing system based on an SMA intelligent material, which solves the problems in the background technology through arrangement.
In order to solve the technical problems, the invention is realized by the following technical scheme:
the invention relates to an airplane deicing system based on an SMA intelligent material, which comprises the following components:
the driving control system is used for designing a scheme of an SMA self-sensing closed-loop control system and realizing reliable, efficient and simple SMA driving control;
the skin structure is designed into a new embedded SMA deicing skin structure, and the theoretical division of the structure is given;
and the self-adaptive deicing rapid reconstruction algorithm is designed to achieve the aims of improving the performance of the reconstruction algorithm and improving the control accuracy.
Further, the resistance-temperature-deformation modeling includes, a Martensite Volume Fraction (MVF) of the SMA material for displaying a crystal phase of martensite and austenite, a resistance and strain of the SMA actuator and other physical quantities depending on a change of the MVF, a displacement (or strain) of the SMA actuator may be estimated by referring to the resistance thereof, a resistance strain behavior of the SMA actuator may be reflected in a self-sensing model in a position control algorithm, and a position control effect of the ladder and sine wave input is verified through experiments to verify effectiveness of the control algorithm of the proposed self-sensing model;
further, the design of the multi-pulse driving-feedback scheme comprises pulse width modulation, wherein the pulse width modulation is a common driving mode, the average power transmitted by an electric signal is reduced by dispersing an effective electric signal into a discrete form, the waveform of corresponding amplitude and frequency required to be synthesized can be equivalently obtained by changing the time width of a pulse according to an area equivalent rule, the SMA can be electrically heated and actuated by utilizing a PWM technology, meanwhile, according to the self-perception characteristic of an SMA material, tiny pulses which are fixed and have a pulse width extremely narrow so as not to play a role in heating and actuating the SMA can be used as excitation signals for measuring the SMA resistance in order to avoid interference between pulse gaps of heating PWM waveforms;
further, the closed-loop control algorithm design includes a position control algorithm for the SMA actuator including a classical proportional-integral-derivative (PID) controller, fuzzy logic, and a self-sensing model, the output of the PID controller determining Vout (t) applied to the SMA actuator, the output being used to electrically heat drive the SMA actuator. It is given by the following formula:
wherein e (t) represents the target deformation ε t (t) and self-sensing shapeChanging epsilon s (t) deviations between; k (K) P Representing a scaling factor; k (K) i Representing an integral coefficient; k (K) d Representing the differential coefficient, the PID parameters are automatically set by the fuzzy logic rules by reference to e (t) and its derivative (Δe (t)/Δt) to minimize disturbances caused by the nonlinear characteristics of the SMA actuator;
further, the skin structure adopts an embedded SMA deicing skin structure, an SMA actuator is embedded into a cavity at the rear end of the front edge of the wing skin, an SMA actuator support is arranged on a rigid structure of the aircraft wing and used for fixing the rear end of the wing, the front end of the SMA actuator is connected with a structure connected with the surface of the skin by willow, when deicing is required, a controller drives the SMA actuator to deform through a cable to generate stress effect, the surface of the skin is broken after ice is attached to the surface and separated from the skin under aerodynamic force effect, an SMA driving element is a pre-strained 3% -5%, ti-50.26% Ni (atomic fraction) alloy wire with the diameter of 1mm, the stress variation range is 250-500 MPa, the stress range meets the requirements of aircraft deicing deformation, and for the arrangement mode and density of stress trigger points, analysis and research are carried out according to the integral structure of the wing and stress deformation dimensions of all parts, and in addition, the power consumption requirement is considered;
further, the self-adaptive deicing rapid reconstruction algorithm comprises the influence of parameters with larger influence such as stress trigger point deformation positions, pulse sequences and the like on deicing effects, the strain deformation positions are controlled through the positions of the stress trigger points, the strain intensity is controlled through controlling driving pulse width, and deicing effect prediction is to establish a set of evaluation criteria for predicting deicing according to the dynamic characteristics of structural deformation and vibration;
further, the deformation position and the pulse sequence comprise a front airfoil surface of a windward part and an icing region of the wing, the curved surface of the current airfoil surface is horizontally unfolded, the upper airfoil surface and the lower airfoil surface of the front airfoil surface respectively correspond to A, B two regions, according to test conditions, the A region and the B region respectively have a stress trigger point, the installation position of the stress trigger point is vertically symmetrical relative to the center line of the front edge of the wing, the action range of the stress trigger point is a small excitation circle, after the front airfoil surface is unfolded, the distance from the center of the excitation circle to the center line of the front edge of the wing is E, the deformation position is changed by changing the installation position of the stress trigger point, and therefore, the deformation position is represented by introducing a parameter gamma:
wherein L is the width of a A, B area behind the front wing surface, E is the larger the distance from the center of the excitation circle to the center line, the farther the deformation position is from the front edge of the wing, the rationality of the wing structure and the installation position of the electric pulse system is considered, the installation position of the SMA actuator is ensured not to be too close to the front edge or the edge zone, and the value is 0.2 to 0.8;
further, the control circuit system can control the pulse sequence of the vibration exciter in the A, B area, so that a pulse interval parameter alpha is introduced to represent the interval between the pulses in the A area and the B area, namely, after the pulse in the A area is started, the pulse in the B area is started after the interval time alpha, the larger the alpha is, the larger the pulse interval between A, B is, and the value range is Oms to 1ms according to actual requirements;
when α= Oms, the A, B region is simultaneously deformed by force, and belongs to double-pulse synchronous impact:
when alpha=0.5 ms, the area A starts to be stressed and deformed at the moment, and the area B starts to be stressed and deformed at intervals of 0.5ms, so that the method belongs to double-pulse step impact;
the invention has the following beneficial effects:
according to the invention, on the basis of setting the SMA self-sensing closed-loop control technology, an SMA deicing technology study based on an embedded SMA deicing skin structure and an SMA driving control technology is carried out, so that the efficiency of an aircraft deicing system is improved, the power consumption is reduced, and meanwhile, the size and the weight of the deicing system are reduced, so that the intelligent aircraft deicing system is suitable for wide popularization and application.
Of course, it is not necessary for any one product to practice the invention to achieve all of the advantages set forth above at the same time.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings that are needed for the description of the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.
FIG. 1 is a block diagram of a system flow of the present invention;
FIG. 2 is a graph of the resistive strain relationship of an SMA actuator of the present invention under heating and cooling cycles;
FIG. 3 is a flow chart of the self-sensing model evaluation strain in accordance with the present invention;
FIG. 4 illustrates an exemplary circuit and pulse waveform of a multi-pulse drive-feedback scheme of the present invention;
FIG. 5 is a block diagram of SMA actuator position control of the present invention;
FIG. 6 is a schematic diagram of an embedded SMA deicing skin structure of the present invention;
FIG. 7 is a schematic diagram of the stress trigger point range of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention.
Example 1
Referring to fig. 1-7, the present embodiment is an aircraft deicing system based on SMA smart materials, comprising:
the driving control system is used for designing a scheme of an SMA self-sensing closed-loop control system and realizing reliable, efficient and simple SMA driving control;
the skin structure is designed into a new embedded SMA deicing skin structure, and the theoretical division of the structure is given;
and the self-adaptive deicing rapid reconstruction algorithm is designed to achieve the aims of improving the performance of the reconstruction algorithm and improving the control accuracy.
The SMA self-sensing closed-loop control system comprises resistance-temperature-deformation modeling, multi-pulse driving-feedback scheme design and closed-loop control algorithm design as shown in figure 1.
In particular, shape Memory Alloys (SMAs) have a Shape Memory Effect (SME) that changes phase between martensite and austenite depending on temperature conditions, most SMA actuator driven applications require a control system, for example, closed Loop Control (CLC) systems driven by SMA actuators require feedback signals (such as displacement, temperature, force, etc.), which typically measure various sensors including hall sensors, tilt-pin sensors, linear Variable Differential Transformers (LVDTs), laser Displacement Sensors (LDS), and strain gauges. However, SMA actuation systems are difficult to miniaturize and when additional sensors are used, obtaining feedback signals becomes very complex. This problem is critical in controlling the SMA drive system, and the greatest advantage of self-sensing feedback is that the feedback signal can be obtained without the need for additional sensors
Wherein as shown in fig. 2 and 3, the resistance-temperature-deformation modeling includes, a Martensitic Volume Fraction (MVF) of the SMA material, the Martensitic Volume Fraction (MVF) of the SMA material being used to display a crystalline phase of martensite and austenite, a physical quantity such as resistance and strain of the SMA actuator being dependent on a change in MVF, a displacement (or strain) of the SMA actuator being estimated by referring to its resistance, a resistive strain behavior of the SMA actuator being reflected in a self-sensing model in a position control algorithm, and a position control effect of the step and sine wave inputs being verified experimentally to verify the effectiveness of the control algorithm of the proposed self-sensing model.
In particular, due to the hysteresis gap caused by the rhombohedral phase (R-phase), in which the heating and cooling paths do not coincide with each other, the R-phase is caused by a decrease in cooling rate and can be observed in SMA actuators at low load levels, it is difficult to accurately estimate strain by referring to the resistance due to the hysteresis gap when the load conditions of the SMA actuators change. In order to achieve accurate position control of the SMA actuator, the effect of the hysteresis gap should be minimized.
As shown in fig. 4, the design of the multi-pulse driving-feedback scheme includes pulse width modulation, which is a common driving mode, by dispersing an effective electric signal into a discrete form so as to reduce the average power transmitted by the electric signal, according to the area equivalent rule, the waveform of the corresponding amplitude and frequency required to be synthesized can be equivalently obtained by changing the time width of the pulse, the SMA can be electrically heated and actuated by using the PWM technology, and meanwhile, according to the self-perception characteristic of the SMA material, tiny pulses can be inserted into the pulse gap of the heating PWM waveform to be used as excitation signals for measuring the SMA resistance, and the tiny pulse width is fixed and the pulse width is extremely narrow so as not to play a role in heating and actuating the SMA.
Wherein the closed-loop control algorithm design includes a position control algorithm for the SMA actuator including a classical proportional-integral-derivative (PID) controller, fuzzy logic, and a self-sensing model, the output of the PID controller determining Vout (t) applied to the SMA actuator, the output being used to electrically heat the SMA actuator. It is given by the following formula:
wherein e (t) represents the deviation between the target deformation ε_t (t) and the self-perceived deformation ε_s (t); k (K) P Representing a scaling factor; k (K) i Representing an integral coefficient; k (K) d Representing the differential coefficient, the PID parameters are automatically set by the fuzzy logic rules by reference to e (t) and its derivative (Δe (t)/Δt) to minimize disturbances caused by the nonlinear characteristics of the SMA actuator; .
As shown in fig. 6, the skin structure adopts an embedded SMA deicing skin structure, an SMA actuator is embedded into a cavity at the rear end of the front edge of a wing skin, an SMA actuator support is mounted on a rigid structure of an aircraft wing and used for fixing the rear end of the wing, the front end of the SMA actuator is connected with a structure connected with the surface of the skin by a willow, when deicing is required, a controller drives the SMA actuator to deform through a cable to generate stress effect, ice attached to the surface of the skin is broken after being stressed and separated from the skin under aerodynamic force, an SMA driving element is a pre-strained 3% -5%, a Ti-50.26% ni (atomic fraction) alloy wire with the diameter of 1mm, the stress change range is 250-500 MPa, the stress range meets the requirements of aircraft deicing deformation, and for the arrangement mode and density of stress trigger points, analysis and research are carried out according to the integral structure of the wing and stress deformation dimensions of all parts, and power consumption requirements are considered.
The self-adaptive deicing rapid reconstruction algorithm comprises the influence of parameters with larger influence such as stress trigger point deformation positions, pulse sequences and the like on deicing effects, the strain deformation positions are controlled through the positions of the stress trigger points, the strain intensity is controlled through controlling driving pulse width, and deicing effect prediction is to establish a set of evaluation criteria for predicting deicing according to the dynamic characteristics of structural deformation and vibration.
As shown in fig. 7, the deformation position and pulse sequence includes a front airfoil surface of a windward part and an icing region of the wing, the curved surface of the current airfoil surface is horizontally unfolded, the upper airfoil surface and the lower airfoil surface of the front airfoil surface respectively correspond to A, B two regions, according to test conditions, the region a and the region B respectively have a stress trigger point, the installation positions of the stress trigger points are vertically symmetrical about the center line of the front edge of the wing, the action range of the stress trigger points is a small excitation circle, after the front airfoil surface is unfolded, the distance from the center of the excitation circle to the center line of the front edge of the wing is set to be E, the deformation position is changed by changing the installation positions of the stress trigger points, and therefore, the parameter gamma is introduced to represent the deformation position:
wherein L is the width of a A, B area behind the front wing surface, E is the larger the distance from the center of the excitation circle to the center line, the farther the deformation position is from the front edge of the wing, the rationality of the wing structure and the installation position of the electric pulse system is considered, the installation position of the SMA actuator is ensured not to be too close to the front edge or the edge zone, and the value is 0.2 to 0.8.
Specifically, in FIG. 7, the upper half A is the upper airfoil, the lower half B is the lower airfoil, and the centerline is the leading edge of the airfoil
The pulse sequence of the vibration exciter in the A, B area can be controlled through a control circuit system, so that a pulse interval parameter alpha is introduced to represent the interval between the pulses in the A area and the B area, namely after the pulse in the A area is started, the pulse in the B area is started after an interval time alpha, the larger the alpha is, the larger the pulse interval between A, B is, and the value range is Oms to 1ms according to actual requirements;
when α= Oms, the A, B region is simultaneously deformed by force, and belongs to double-pulse synchronous impact:
when alpha=0.5 ms, the A area starts to be stressed and deformed at the moment, and the B area starts to be stressed and deformed at intervals of 0.5ms, and the method belongs to double-pulse step impact.
In the description of the present specification, a description referring to terms "one embodiment," "some embodiments," "examples," "specific examples," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment or example is included in at least one embodiment or example of the present invention. In this specification, schematic representations of the above terms are not necessarily directed to the same embodiment or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples. Furthermore, the different embodiments or examples described in this specification and the features of the different embodiments or examples may be combined and combined by those skilled in the art without contradiction.
In the description of the present invention, it should also be noted that, unless explicitly specified and limited otherwise, the terms "disposed," "mounted," "connected," and "connected" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention can be understood by those of ordinary skill in the art according to the specific circumstances.
Finally, it should be noted that: the foregoing description is only a preferred embodiment of the present invention, and the present invention is not limited thereto, but it is to be understood that modifications and equivalents of some of the technical features described in the foregoing embodiments may be made by those skilled in the art, although the present invention has been described in detail with reference to the foregoing embodiments. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (9)
1. An aircraft deicing system based on SMA intelligent materials is characterized in that; comprising the following steps:
the driving control system is used for designing a scheme of an SMA self-sensing closed-loop control system and realizing reliable, efficient and simple SMA driving control;
the skin structure is designed into a new embedded SMA deicing skin structure, and the theoretical division of the structure is given;
and the self-adaptive deicing rapid reconstruction algorithm is designed to achieve the aims of improving the performance of the reconstruction algorithm and improving the control accuracy.
2. An SMA smart material based aircraft deicing system of claim 1, wherein said SMA self-aware closed-loop control system comprises resistance-temperature-deformation modeling, multi-pulse drive-feedback scheme design, closed-loop control algorithm design.
3. An aircraft deicing system based on SMA smart material according to claim 2, wherein said resistance-temperature-deformation modeling comprises SMA material Martensite Volume Fraction (MVF) for displaying martensite and austenite crystallographic phases, the physical quantities of resistance and strain of the SMA actuator being dependent on the change in MVF, the displacement (or strain) of the SMA actuator being estimated by reference to its resistance, the resistive strain behavior of the SMA actuator being reflected in a self-sensing model in a position control algorithm, the position control effect of the ladder and sine wave inputs being experimentally verified to verify the effectiveness of the control algorithm of the proposed self-sensing model.
4. An aircraft deicing system based on SMA intelligent materials according to claim 2, characterized in that said design of the multi-pulse drive-feedback scheme comprises pulse width modulation, which is a common driving method, by dispersing the effective electric signal into discrete form to reduce the average power delivered by the electric signal, the equivalent rule of area can be obtained by changing the time width of the pulse to equivalently obtain the waveform of the corresponding amplitude and frequency required to be synthesized, the SMA can be electrically heated and actuated by PWM technique, while for the self-perception feature of SMA material, the insertion of micro pulse as the excitation signal of SMA resistance measurement in the pulse gap of the heating PWM waveform is fixed and the pulse width is so narrow as not to exert heating and actuating effect on SMA.
5. An aircraft deicing system based on SMA smart material according to claim 2, wherein said closed-loop control algorithm design comprises a position control algorithm for SMA actuators comprising a classical proportional-integral-derivative (PID) controller, fuzzy logic, and a self-sensing model, the output of the PID controller determining Vout (t) applied to the SMA actuator, the output being used to electrically heat drive the SMA actuator. It is given by the following formula:
wherein e (t) represents a deviation between the target deformation and the self-sensed deformation; KP represents a scaling factor; ki represents an integral coefficient; kd represents the differential coefficient and the PID parameters are automatically set by the fuzzy logic rules by reference to e (t) and its derivative (Δe (1)) to minimize disturbances caused by the nonlinear characteristics of the SMA actuator.
6. The aircraft deicing system based on the SMA intelligent material according to claim 2, wherein the skin structure adopts an embedded SMA deicing skin structure, an SMA actuator is embedded into a cavity at the rear end of the front edge of the wing skin, an SMA actuator support is arranged on a rigid structure of the aircraft wing for fixing the rear end of the anchor, the front end of the SMA actuator is connected with a structure which is connected with the surface of the skin by willow, when deicing is required, a controller drives the SMA actuator to deform through a cable to generate stress, the ice attached to the surface of the skin breaks after being stressed and is separated from the skin under the action of aerodynamic force, the SMA driving element adopted is a pre-strained 3% -5%, ti-50.26% Ni (atomic fraction) alloy wire with the diameter of 1mm, the stress change range is 250-500 MPa, the stress change range meets the requirements of the aircraft deicing deformation, the arrangement mode and the density of stress trigger points are analyzed and studied according to the integral structure of the wing and the stress deformation dimensions of each part, and the power consumption requirements are considered.
7. The aircraft deicing system based on SMA intelligent material of claim 1, wherein the adaptive deicing rapid reconstruction algorithm comprises the influence of stress trigger point deformation position, pulse sequence and other parameters with larger influence on deicing effect, the strain force deformation position is controlled by the position of the stress trigger point, the strain force intensity is controlled by controlling the driving pulse width, and deicing effect prediction is to establish a set of evaluation criteria for predicting deicing according to the dynamics characteristics of structural deformation and vibration.
8. The system of claim 7, wherein the deformation position and pulse sequence comprises a front airfoil surface of a windward part and an icing area of the wing, the curved surface of the front airfoil surface is horizontally unfolded, the upper airfoil surface and the lower airfoil surface of the front airfoil surface respectively correspond to A, B two areas, according to the test condition, the area A and the area B respectively have a stress trigger point, the installation positions of the stress trigger points are vertically symmetrical about the center line of the front edge of the wing, the action range of the stress trigger points is a small excitation circle, after the front airfoil surface is unfolded, the distance from the center of the excitation circle to the center line of the front edge of the wing is E, the deformation position is changed by changing the installation positions of the stress trigger points, and therefore, the parameter gamma is introduced to represent the deformation position:
wherein L is the width of a A, B area behind the front wing surface, E is the larger the distance from the center of the excitation circle to the center line, the farther the deformation position is from the front edge of the wing, the rationality of the wing structure and the installation position of the electric pulse system is considered, the installation position of the SMA actuator is ensured not to be too close to the front edge or the edge zone, and the value is 0.2 to 0.8.
9. An aircraft deicing system based on SMA intelligent material according to claim 7, wherein the pulse sequence of the exciter of region A, B is controlled by control circuitry, so that pulse interval parameter α is introduced to represent the interval between the pulses of regions a and B, i.e. after the start of the pulse of region a, after the interval time α, the pulse of region B starts, the greater α, the greater the pulse interval between A, B, the range of values being Oms to 1ms, according to practical requirements;
when α= Oms, the A, B region is simultaneously deformed by force, and belongs to double-pulse synchronous impact:
when alpha=0.5 ms, the A area starts to be stressed and deformed at the moment, and the B area starts to be stressed and deformed at intervals of 0.5ms, and the method belongs to double-pulse step impact.
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| CN202310371323.0A CN116374177A (en) | 2023-04-07 | 2023-04-07 | Airplane deicing system based on SMA intelligent material |
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4399967A (en) * | 1980-12-09 | 1983-08-23 | Lockheed Corporation | Staggered coil and nose-torquer electromagnetic pulse deicing systems |
| US5686003A (en) * | 1994-06-06 | 1997-11-11 | Innovative Dynamics, Inc. | Shape memory alloy de-icing technology |
| US20080075593A1 (en) * | 2006-05-17 | 2008-03-27 | Simon Read | Apparatus for preventing ice accretion |
| CN204368433U (en) * | 2014-12-23 | 2015-06-03 | 中国民航大学 | A kind of wing deicing device driven based on SMA |
| CN112055684A (en) * | 2018-03-29 | 2020-12-08 | Gkn航空服务有限公司 | Deicing system |
-
2023
- 2023-04-07 CN CN202310371323.0A patent/CN116374177A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4399967A (en) * | 1980-12-09 | 1983-08-23 | Lockheed Corporation | Staggered coil and nose-torquer electromagnetic pulse deicing systems |
| US5686003A (en) * | 1994-06-06 | 1997-11-11 | Innovative Dynamics, Inc. | Shape memory alloy de-icing technology |
| US20080075593A1 (en) * | 2006-05-17 | 2008-03-27 | Simon Read | Apparatus for preventing ice accretion |
| CN204368433U (en) * | 2014-12-23 | 2015-06-03 | 中国民航大学 | A kind of wing deicing device driven based on SMA |
| CN112055684A (en) * | 2018-03-29 | 2020-12-08 | Gkn航空服务有限公司 | Deicing system |
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