Patent Publication Number: US-2021188450-A1

Title: Electric field sensor with sensitivity-attenuating ground ring

Description:
RELATED APPLICATIONS 
     This patent application is a continuation of U.S. patent application Ser. No. 15/571,588, filed Nov. 3, 2017, entitled “Apparatus and Method of Monitoring for In-Flight Aircraft Engine Ice Crystal Accretion”, which is a § 371 of International Patent Application Serial No. PCT/CA2016/050517, filed May 5, 2016, which claims the benefit of the filing date of U.S. Patent Application Ser. No. 62/156,891, filed May 5, 2015, the contents of all of which are incorporated herein by reference. 
    
    
     FIELD 
     This patent application relates to an electric field sensor. In particular, this patent application relates to an electric field sensor that may be used to monitor for the in-flight accretion of ice crystals on aircraft engine surfaces. 
     BACKGROUND 
     Aircraft may be exposed to various atmospheric icing conditions in-flight, such as small water droplets less than 25 um diameter (e.g. clouds, fog) and super-cooled large water droplets greater than 25 um diameter (e.g. freezing rain, freezing drizzle). Ice accretion resulting from exposure to these atmospheric icing conditions, in and around the warm surfaces of an aircraft engine, such as the cowling and/or air intake ducts of turbofan engines, the carburetor mouth and/or throttle body of piston engines, and even the internal surfaces of the engine, can cause engine damage and/or a sudden loss in engine output power and aircraft stability. Therefore, many modern aircraft engines incorporate countermeasures, such as heating systems, to reduce the likelihood or extent of engine ice accretion on the external and internal engine surfaces. However, over the past 20 years there have been over 100 reported cases of aircraft engine power loss as a result of ice accretion on engine surfaces. 
     SUMMARY 
     This patent application describes an electric field sensor that includes an insulating substrate, a plurality of electrodes, an insulator, a plurality vias, and a ground ring. 
     The electrodes are disposed on the substrate. The insulator is disposed over the electrodes. 
     The vias are coupled to the electrodes and extend through the substrate at a right angle to the electrodes. 
     The ground ring is disposed around the electrodes and the vias and is configured to attenuate a sensitivity of the sensor to electric fields outwards of the ground ring. 
     The plurality of electrodes may include a plurality of first electrodes and a plurality of second electrodes that are interleaved and non-contacting with the plurality of first electrodes. The first electrodes are not in contact with each other, the second electrodes are not in contact with each other, and the plurality of first electrodes are disposed parallel to the plurality of second electrodes on the substrate. 
     The plurality of vias may include a first via portion and a second via portion. The vias of the first via portion are coupled to the plurality of first electrodes, and the vias of the second via portion are coupled to the plurality of second electrodes. 
     The insulator may be configured to be mounted substantially flush with the surface of an aircraft engine. 
     The insulator may be aerodynamically-matched to the surface of the aircraft engine. 
     The insulator may be thermally-matched to the surface of the aircraft engine. 
     Proposed aircraft engine surfaces include the cowling of a turbofan engine, the air intake duct of a turbofan engine, the carburetor mouth of a piston engine, the throttle body of a piston engine, and an internal engine surface. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The foregoing aspects will now be described, by way of example only, with reference to the accompanying drawings, in which: 
         FIG. 1  is a schematic view of a matter accumulation monitoring system, depicting the electric field sensor unit and a sensor processing unit; 
         FIG. 2 a    is a perspective view of the electric field sensor; 
         FIG. 2 b    is a top plan view of the electric field sensor; 
         FIGS. 2 c  and 2 d    are transverse cross-sectional views of the electric field sensor; 
         FIG. 3  is a schematic view of the structure of the sensor processing unit; and 
         FIG. 4  is a flow-chart depicting a method of operation of the matter accumulation monitoring system. 
     
    
    
     DETAILED DESCRIPTION 
     1. Matter Accumulation Monitoring System: Overview 
     Turning now to  FIG. 1 , there is shown a matter accumulation monitoring system, denoted generally as  100 , comprising one or more electric field sensor units  200 , a sensor processing unit (SPU)  300 , and a wiring harness  102  interconnecting the electric field sensor units  200  and the sensor processing unit (SPU)  300 . 
     As shown, the electric field sensor units  200  are preferably separate and distinct from the SPU  300 . Although the functionality of the SPU  300  may also be incorporated into each of the electric field sensor units  200 , it should be understood that maintaining the electric field sensor units  200  separate and distinct from the SPU  300  reduces the physical size of the electric field sensor units  200  and allows the electric field sensor units  200  to be better aerodynamically-matched and thermally-matched to the associated aircraft surfaces/engine surfaces. 
     2. Electric Field Sensor Unit 
     Each electric field sensor unit  200  is disposed proximate a surface of an aircraft, at a respective region thereof, to monitor the accumulation of matter on the respective region of the aircraft surface while the aircraft is in-flight. 
     The matter to be monitored may be ice, and one or more of the electric field sensor units  200  may be mounted on and thermally-coupled to an exterior surface of an aircraft at regions thereof that are prone to the accumulation of ice while the aircraft is in-flight. Preferably, the electric field sensor units  200  are aerodynamically-matched to the aircraft surfaces being monitored so as to not disturb the normal airflow and hence the normal buildup of ice on the associated aircraft surfaces. Further, preferably the electric field sensor units  200  are also thermally-matched to the associated aircraft surfaces so as to again not interfere with the normal buildup of ice on the aircraft surfaces. 
     More preferably, one or more of the electric field sensor units  200  is mounted to and thermally-coupled to an engine surface of an aircraft at regions of the engine that are prone to the accretion of ice crystals while the aircraft is in-flight. Typically an aircraft is fitted with a plurality of the electric field sensor units  200 . It is expected that, through experimentation, the person of ordinary skill could determine the appropriate placement of the electric field sensor units  200  to detect ice crystal accretion on engine surfaces while the aircraft is in flight. 
     Preferably, the electric field sensor units  200  are aerodynamically-matched to the engine surfaces being monitored so as to not disturb the normal airflow and hence the normal buildup of ice crystals on the engine surfaces. Accordingly, the electric field sensor units  200  may be flush-mounted with the respective engine surfaces. As non-limiting examples, one or more of the electric field sensor units  200  may be flush-mounted with a cowling of a turbofan engine, an air intake duct of a turbofan engine, a carburetor mouth of a piston engine, a throttle body of a piston engine, or an internal engine surface. 
     Although conventional ice mitigation countermeasures have been successful in limiting ice accretion on engine surfaces from exposure to small water droplets less than 25 um diameter (e.g. clouds, fog) and super-cooled large water droplets greater than 25 um diameter (e.g. freezing rain, freezing drizzle), the inventors have learned that fully glaciated water (e.g. ice crystals) can accumulate on the heated surfaces of aircraft engines. In contrast to an aircraft wing where the wing surface remains cold while the aircraft is in flight, the inventors believe that heat transfer from an engine surface while the aircraft is in flight initially causes the ice crystals to melt, but that rapid air flow over the engine surface subsequently cools the melt and causes the ice crystals to adhere to the engine surface. Therefore, so as to again not interfere with the normal buildup of ice on the engine surfaces, preferably the electric field sensor units  200  are also thermally-matched to the associated engine surfaces in the sense that the field sensor units  200  absorb heat from (and introduce heat into) the airstream passing over the electric field sensor units  200  at the same rate as the associated aircraft surface/engine surfaces. 
     As shown in  FIG. 1 , preferably each electric field sensor unit  200  includes a signal amplifier  204 , and an electric field sensor  202  coupled to the signal amplifier  204 . Optionally, the electric field sensor unit  200  may also include a digital temperature sensor  203 , and a heater (not shown). 
     Each electric field sensor  202  generates a time-varying local electric field at the engine surface, and a current is produced in the electric field sensor  202  resulting from the time-varying local electric field. The magnitude and phase of the resulting current varies with the characteristics of the material that is within the electric field established by the electric field sensor  202 . Accordingly, the electric field sensor(s)  202  together provide periodic data samples indicative of the accretion of ice crystals on the associated aircraft surface/engine surface. 
     The temperature sensors  203  monitor the temperature of the associated aircraft surfaces/engine surfaces and are typically used to calibrate the electric field sensor  202 . Alternately, or additionally, the temperature sensor(s)  203  may be used in the detection of ice crystals on the associated aircraft surfaces/engine surfaces. 
     As shown in  FIGS. 2 a , 2 b  and 2 c   , each electric field sensor  202  comprises an insulating substrate  206 , a plurality of electrodes  208  disposed on the substrate  206 , and a plurality of vias  210  coupled to the plurality of electrodes  208  and extending downwardly through the substrate  206 . Preferably, the substrate  206  comprises a non-conductive material, such as ceramic, although other non-conductive materials may be used. The electrodes  208  are typically formed on the substrate  206  using conventional printed circuit board or integrated circuit manufacturing techniques. The electrodes  208  extend across the top surface of the substrate  206  in a substantially parallel fashion, such that the electrodes  208  do not contact one another on the top surface of the substrate  206 . 
     The electrodes  208  are segregated into a first electrode portion  208   a , and a second electrode portion  208   b . The electrodes  208  of the first electrode portion  208   a  are not in contact with one another and extend from one end  212   a  of the substrate  206 , and the electrodes  208  of the second electrode portion  208   b  are not in contact with one another and extend from the opposite end  212   b  of the substrate  206 . The electrodes  208  of the first electrode portion  208   a  are interleaved with and are disposed substantially parallel to the electrodes  208  of the second electrode portion  208   b  in the centre region  214  of the top surface of the substrate  206 . 
     Typically, each via  210  comprises a plated through-hole extending from one end of a respective electrode  208 , through the substrate  206 , towards the bottom surface  216  of the electric field sensor unit  200 . Alternately, the vias  210  may be provided as conductive traces or wires extending in a similar manner. The vias  210  are segregated into a first via portion  210   a  and a second via portion  210   b . The vias  210  of the first via portion  210   a  are coupled to the first electrode portion  208   a , and the vias  210  of the second via portion  210   b  are coupled to the second electrode portion  208   b.    
     Each via  210  is connected to a respective electrode  208  adjacent the respective end  212 , and extends at a right angle from the electrode  208  through the substrate  206 , from the top surface of the electric field sensor unit  200  towards the bottom surface  216  of the electric field sensor unit  200 . Preferably, the electric field sensor  202  also includes a ground ring  218 , disposed on the substrate  206  laterally outwards from the electrodes  208  and the vias  210 . The ground ring  218  may be connected to electrical grounds of the signal amplifier  204 . With this configuration, the sensitivity of the electric field sensor  202  to electric fields outside the centre region  214  is less than conventional electric field sensors. 
     Each via  210  of the first via portion  210   a  terminates at a first conductive plate  220  that is embedded within the substrate  206  of the electric field sensor unit  200 . Similarly, each via  210  of the second via portion  210   b  terminates at a second conductive plate (not shown) that is embedded within the substrate  206  (but separate from the first conductive plate  220 ). 
     The electric field sensor  202  preferably also includes a top electrically-insulating layer  222  disposed over the electrodes  208 , and a bottom electrically-insulating layer  224  disposed over the bottom surface  216  of the electric field sensor unit  200 . As discussed, preferably the electric field sensor units  200  are aerodynamically-matched to the aircraft surfaces/engine surfaces being monitored. Accordingly, preferably the top insulating layer  222  is aerodynamically-matched to the shape and contour of the associated aircraft surface/engine surface so as to not disturb the normal airflow and hence the normal buildup of ice crystals on the aircraft surface/engine surface. Typically, the top insulating layer  222  is disposed substantially flush with the associated aircraft surface/engine surface. As discussed, suitable engine surfaces include, but are not limited to, the cowling of a turbofan engine, the air intake duct of a turbofan engine, the carburetor mouth of a piston engine, the throttle body of a piston engine, and an internal engine surface. 
     As discussed, preferably the electric field sensor units  200  are thermally-coupled and thermally-matched to the aircraft surfaces/engine surfaces being monitored. Accordingly, so as to further not interfere with the normal buildup of ice on the aircraft surface/engine surface, preferably the top insulating layer  222  is thermally-matched to the associated aircraft surface/engine surface such that the top insulating layer  222  absorbs heat from (and introduces heat into) the airstream passing over the insulating layer at the same rate as the associated aircraft surface/engine surface. Typically, the insulating layers  222 ,  224  each comprises a ceramic or glass, although the ceramic alumina is preferred due to its hardness and thermal conductivity 
     Preferably, the signal amplifier  204  is incorporated into the electric field sensor unit  200  below the electric field sensor  202  and is connected to the vias  210  of the electric field sensor  202  via the conductive plates  220 . First and second electrically-conductive pins  226   a ,  226   b  extend from the signal amplifier  204  and out the bottom surface  216  of the electric field sensor unit  200 . 
     As will be explained, the sensor processing unit (SPU)  300  applies time-varying analog voltage signals to the first and second conductive pins  226   a ,  226   b  via the wiring harness  102 . The signal amplifier  204  receives the time-varying analog voltage signals via the conductive pins  226   a ,  226   b , amplifies the analog voltage signals, and applies the amplified analog voltage signals to the electrodes  208  of the electric field sensor  202  via the vias  210 . Preferably, the analog voltage signal applied to the first electrode portion  208   a  is complementary (i.e. 180 degrees out of phase) with the analog voltage signal applied to the second electrode portion  208   b.    
     The amplified analog voltage signals cause the electrodes  208  to apply a time-varying electric field to the aircraft surface/engine surface that is proximate the electric field sensor  202 . The time-varying electric field causes an analog current to be produced in the electrodes  208  of the electric field sensor  202 . The signal amplifier  204  detects the resulting analog current via the vias  210  of the electric field sensor  202 , generates an analog voltage signal from the resulting current, and applies the generated analog voltage signal to the first and second conductive pins  226   a ,  226   b . The SPU  300  detects the generated analog voltage signal at the first and second conductive pins  226   a ,  226   b  via the wiring harness  102 . 
     Preferably, the temperature sensor  203  is incorporated into the associated electric field sensor  202  proximate the top surface of the substrate  206 , and the electric field sensor  202  includes a via  210  that extends through the substrate  206  from proximate the top surface towards the bottom surface  216  of the electric field sensor unit  200 . A third electrically-conductive pin  222   c  extends from the via  210  of the temperature sensor  203  and out the bottom surface  226  of the electric field sensor unit  200 . Alternately, each temperature sensor  203  may be disposed in a substrate that is separate from the substrate  206  of the electric field sensor unit  200 . 
     The temperature sensor  203  generates a serial digital output signal that includes temperature measurement samples of the ambient temperature proximate the electric field sensor  202 , and applies the generated digital temperature measurement signal to the third conductive pin  226   c . The SPU  300  receives the temperature measurement samples, that are output at the third conductive pin  226   c , via the wiring harness  102 . 
     Preferably, the heater is also incorporated into the associated electric field sensor  202  proximate the top surface of the substrate  206 , and is connected to the SPU  300  via the wiring harness  102 . As will be explained, the SPU  300  uses the heater to melt ice that may have accumulated proximate an electric field sensor unit  200 . However, so as to avoid frustrating the detection of ice at an electric field sensor unit  200  (by interfering with the thermal matching between the electric field sensor unit  200  and the aircraft surfaces/engine surfaces being monitored), the heater is otherwise typically inactive while the aircraft is in flight. 
     3. Sensor Processing Unit 
     The sensor processing unit (SPU)  300  is typically disposed within the aircraft cockpit, and is coupled to the electric field sensor units  200  via the wiring harness  102 . As shown in  FIG. 3 , preferably the SPU  300  includes a sensor monitor  302 , and an analog signal generator  320 , and an A/D converter  322 . 
     The sensor monitor  302  comprises a non-volatile non-transient memory  304 , a monitor interface  306 , and a central processing unit (CPU)  308  that is coupled to the non-volatile memory  304  and the monitor interface  306 . The monitor interface  306  may interface the SPU  300  with the instrumentation of the aircraft cockpit to thereby provide pilots with a substantially-real time assessment of the accumulation of matter on the aircraft surfaces/engine surfaces. 
     The non-transient memory  304  may be provided as an electronic memory, a magnetic disc and/or an optical disc, and may include a signatures database  350  of one or more predetermined matter accumulation profiles. Each predetermined matter accumulation profile is associated with a particular characteristic (e.g. thickness, matter type) of the matter accumulated, and comprises a corresponding time-series of complex admittance values and optionally temperature values. Typically each predetermined matter accumulation profile includes a time-series of complex admittance values and optionally temperature values for a particular matter accumulation characteristic while the aircraft is in flight. The time-series of complex admittances and temperatures (if included) in each matter accumulation profile may be predetermined experimentally and/or via computer modelling, and may be typically stored in the memory  304  prior to installation of the SPU  300  in the aircraft. 
     The non-transient memory  304  also stores processing instructions for the SPU  300  which, when executed by the CPU  308 , may define a signature monitor  310  and a signal processor  330 . The signature monitor  310  commands the analog signal generator  320  to generate and apply a time-varying analog voltage signal to the signal amplifier  204  of each of the electric field sensor units  200 . As discussed, the signal amplifier  204  amplifies the received analog voltage signal, and applies the amplified analog voltage signals to the electrodes  208  of the electric field sensor  202 . The amplified analog voltage signal cause the electrodes  208  to apply a time-varying electric field to the aircraft surface/engine surface that is proximate the electric field sensor  202 . The time-varying electric field causes an analog current to be produced in the electrodes  208  of the electric field sensor  202 . The signal amplifier  204  detects the resulting analog current, and generates an analog voltage signal from the resulting current. 
     The A/D converter  322  periodically digitizes, over a measurement time span, the analog voltage signals generated by the signal amplifier  204  of the electric field sensor(s)  202  and provides the signal processor  330 , in substantially real-time, with the digitized version of the analog voltage signals (hereinafter the “digitized current measurement samples”). 
     The signal processor  330  uses the digitized current measurement samples from the A/D converter  322 , and optionally the digital temperature measurement samples from the temperature sensors  203 , to create a time-series of measurement data sets. Each measurement data set includes a magnitude measurement and a phase measurement. Preferably, the signal processor  330  derives the magnitude and phase measurements from the digitized current measurement samples by referencing the magnitude and phase of the current produced in the electric field sensor(s)  202  respectively to the magnitude and phase of the applied sensor voltage signals. In effect, then, the magnitude and phase measurements are complex admittance measurements. However, for ease of reference, the current magnitude and current phase measurements (referenced to the applied sensor voltage) will be referred to hereinafter respectively as magnitude measurements and phase measurements and collectively as complex admittance measurements. 
     As discussed, the temperature sensors  203  are typically used to calibrate the electric field sensor  202 . Therefore, for example, for calibration, the signal processor  330  may use temperature samples received from the temperature sensor(s)  203  to compute weight factors to apply to the complex admittance measurements. 
     In addition to the complex admittance measurements, each measurement data set may optionally also include the digitized temperature measurement sample that was taken when the associated digitized current measurement sample was generated. Each measurement data set may also identify the time at which the complex admittance measurements and temperature measurements were taken. 
     The inventors have determined that the values for the complex admittance measurements, and also the variability (e.g. rate of change, range of fluctuation) in those values between successive measurements, varies with the characteristics of the matter accumulating on the aircraft surface/engine surface. Therefore, to facilitate the detection of matter accumulating on the aircraft surface/engine surface, and the differentiation between different matter accumulating on the aircraft surface/engine surface, the signal processor  330  provides the signature monitor  310  with the time-series of the complex admittance measurements (and optionally temperature measurements). The signature monitor  310  is configured to use the time-series of complex admittance measurements (and optionally temperature measurements) to generate an assessment of the instantaneous accumulation of matter on the aircraft surface/engine surface in substantially-real time. 
     The signature monitor  310  may generate the assessment by querying the signatures database  350  with the received measurement data sets for corresponding (identical or similar) predetermined matter accumulation profiles. Alternately, the signature monitor  310  (or the signal processor  330 ) may be configured with a matter assessment algorithm that generates the assessment of the instantaneous accumulation of matter from the measurement data sets. 
     The signature monitor  310  may then generate and transmit an alarm signal to the aircraft cockpit in real time if one or more of the received measurement data sets correlates with one of the predetermined matter accumulation profiles, indicating that ice crystals have accreted on an aircraft surface/engine surface. Alternately, the signature monitor  310  may generate and transmit an alarm signal to the aircraft cockpit in real time if the matter assessment algorithm determines that ice crystals have accreted on an aircraft surface/engine surface. 
     As discussed above, the signature monitor  310  and the signal processor  330  may each be implemented as a set of computer processing instructions. However, the implementation of the signature monitor  310  and the signal processor  330  is not so limited, but may each be implemented instead in electronics hardware, such as a field programmable logic gate array (FPGA) or a complex programmable logic device (CPLD). 
     4. Matter Accumulation Monitoring System: Method of Operation 
     The method of operation of the matter accumulation monitoring system  100  will now be described with reference to  FIG. 4 . 
     At step S 400 , the matter accumulation monitoring system  100  applies time-varying electric fields to the aircraft surfaces and/or engine surfaces. To do so, the analog signal generator  320  of the SPU  300  may apply a time-varying analog voltage signal to the electrodes  208  of the electric field sensor(s)  202  (via the signal amplifier  204 ), with the voltage signal applied to the first electrode portion  208   a  being 180 degrees out of phase with the voltage signal applied to the second electrode portion  208   b . The alternating electric fields are typically applied to the aircraft surfaces/engine surfaces while the aircraft is in flight. 
     At step S 402 , the SPU  300  generates a time-series of measurement data sets, each measurement data set comprising a complex admittance measurement of the resulting current produced in the electric field sensor unit  200 , and optionally the temperature of the aircraft surfaces/engine surfaces. To do so, the A/D converter  322  may periodically sample the resulting current produced in the associated electric field sensor(s)  202 , and the signal processor  330  may generate the time-series of measurement data sets from the digitized current measurement samples and optionally the temperature measurement samples. As noted above, the values in the measurement data sets, and the variability in those values between successive measurement data sets, will vary in accordance with the characteristics of the matter accumulating on the aircraft surfaces/engine surfaces. 
     At step S 404 , the SPU  300  generates an assessment of the instantaneous accumulation of matter (e.g. ice crystals) on the aircraft surfaces/engine surfaces. To do so, the signature monitor  310  may query the signatures database  350  with the time-series of measurements from one or more of the received measurement data sets, and generate the assessment from a correlation between the time-series of measurements and the predetermined matter accumulation profiles stored in the signatures database  350 . As discussed above, each predetermined matter accumulation profile is associated with a particular characteristic (e.g. thickness, matter type) of accumulated matter (e.g. ice crystals). 
     Where the time-series of received measurements correlates well with a particular matter accumulation profile (the received measurement data sets are identical or similar to one or more of the predetermined matter accumulation profiles), the SPU  300  may generate the assessment from the characteristics of the located matter accumulation profile. 
     However, typically an aircraft will be fitted with a plurality of electric field sensor units  200 , each located at a respective aircraft surface/engine surface. Accordingly, the SPU  300  may generate the assessment by querying the signatures database  350  with the measurement data sets received from a plurality of the electric field sensor units  200 , and generate the assessment from a correlation between the received measurement data sets and a plurality of the predetermined matter accumulation profiles. In this variation, the assessment may include characteristics from the various matter accumulation profiles. Alternately, the signature monitor  310  may generate the assessment by applying the measurement data sets as inputs to a matter assessment algorithm. 
     The SPU  300  may thereafter transmit the results of the assessment to the aircraft cockpit for display on cockpit instrumentation. Alternately, or additionally, the SPU  300  may initiate an automated control action (e.g. activate an alarm, invoke a reduction in the aircraft speed) in accordance with a result of the correlation. For example, the SPU  300  may activate an alarm if the signature monitor  310  determines that ice crystals in excess of a threshold amount have accreted on an aircraft surface/engine surface). 
     After an excess ice accumulation has been detected, the SPU  300  may activate the heater of the associated electric field sensor unit(s)  200  to thereby melt the ice accumulation at the respective electric field sensor unit(s)  200 . During this phase, the SPU  300  may continue to use the measurement data sets received from the electric field sensor units  200  to generate an assessment of the instantaneous accumulation of ice crystals on the aircraft surfaces/engine surfaces, and may maintain the heater(s) active until the assessment indicates that the ice accumulation proximate the respective electric field sensor unit(s)  200  has been shed. Accordingly, the SPU  300  may periodically activate and deactivate the heater(s) as the aircraft is in flight, in accordance with the detected ice accumulation. 
     The SPU  300  may also correlate the measurement data sets received from the electric field sensor units  200  (while a heater is active and the ice accumulation is being shed) with the matter accumulation profiles to further confirm the identity of the substance that had accumulated proximate the electric field sensor unit(s)  200 .