Patent Publication Number: US-2023134390-A1

Title: Air inlet strut for aircraft engine

Description:
TECHNICAL FIELD 
     The application relates generally to aircraft engines and, more particularly, to air inlets for aircraft engines. 
     BACKGROUND 
     It is sometimes desirable to know the pressure at different locations of an aircraft engine, such as at or near the inlet of a compressor of the aircraft engine. A technique for determining this pressure uses ambient atmospheric pressure or an aircraft total pressure (pitot). However, this technique may not capture the effect on the pressure at the inlet of the compressor caused by various operational or installation effects such as losses due to icing, variations in angle of attack, inlet by-pass flow, inertial particle separators, inlet barrier filters, and/or the left/right/center installation of the aircraft engine on the aircraft. 
     SUMMARY 
     There is disclosed an aircraft engine, comprising: an air inlet duct extending from an inlet through which air is configured to enter the aircraft engine to an outlet within the aircraft engine; and at least one strut having a leading edge and a trailing edge and extending across at least part of the air inlet duct, the at least one strut having a strut passage and a plurality of static pressure measurement taps spaced apart on the trailing edge and in fluid communication with the strut passage. 
     There is disclosed an aircraft engine, comprising: an engine casing having an annular wall about a center axis of the aircraft engine, and an air inlet duct extending from an opening in the annular wall to an outlet of the air inlet duct radially inward of the annular wall; and at least one strut extending across the air inlet duct and between a leading edge and a trailing edge, the at least one strut having a strut passage extending within the at least one strut and a plurality of static pressure measurement taps spaced apart on the trailing edge and in fluid communication with the strut passage. 
     There is disclosed a method of obtaining static pressure in an inlet of a compressor of an aircraft engine. The method includes obtaining the static pressure at spaced-apart locations on a trailing edge of a strut extending across the inlet. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG.  1 A  is a schematic cross sectional view of a gas turbine engine; 
         FIG.  1 B  is a schematic cross sectional view of another gas turbine engine; 
         FIG.  1 C  is a schematic cross sectional view of another gas turbine engine; 
         FIG.  2 A  is an enlarged, schematic cross-sectional view of region IIA in  FIG.  1 A ; 
         FIG.  2 B  is another cross-sectional view of region IIA in  FIG.  1 A ; 
         FIG.  3    is a perspective view of a strut shown in  FIG.  2 B ; 
         FIG.  4 A  is a cross-sectional view of region IVA in  FIG.  1 B ; and 
         FIG.  4 B  is an enlarged cross-sectional view of region IVB in  FIG.  4 A . 
     
    
    
     DETAILED DESCRIPTION 
       FIGS.  1 A to  1 C  illustrate different gas turbine engines  10  of a type preferably provided for use in subsonic flight. Each of the gas turbine engines  10  generally comprises in serial flow communication an air inlet  11 , a compressor section  12  for pressurizing the air from the air inlet  11 , a combustor  13  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, a turbine section  14  for extracting energy from the combustion gases, and an exhaust outlet  15  through which the combustion gases exit the gas turbine engine  10 . The gas turbine engine  10  have a longitudinal center axis  17  about which components rotate. In the gas turbine engines  10  shown in  FIGS.  1 A and  1 C , the air inlet  11  is positioned forward of the compressor section  12 , whereas in the gas turbine engine  10  shown in  FIG.  1 B , the air inlet  11  is positioned aft of the compressor section  12 . The gas turbine engine  10  of  FIG.  1 A  includes a driven gear train  16 A mounted at a front end of the gas turbine engine  10 , and is an example of a “turboshaft” gas turbine engine  10 . The gas turbine engine  10  of  FIG.  1 B  includes a propeller  16 B which provides thrust for flight and taxiing, and is an example of a “turboprop” gas turbine engine  10 . The gas turbine engine  10  of  FIG.  1 C  includes a fan  16 C which provides thrust for flight, and is an example of a “turbofan” gas turbine engine  10 . 
     The gas turbine engines  10  (sometimes referred to herein simply as “engines  10 ”) have a central core  18  through which gases flow and which includes some of the turbomachinery of the engine  10 . The engine  10  of  FIG.  1 B  is a “reverse-flow” engine  10  because gases flow through the core  18  from the air inlet  11  at a rear portion, to the exhaust outlet  15  at a front portion. This is in contrast to “through-flow” gas turbine engines  10 , such as those shown in  FIGS.  1 A and  10   , in which gases flow through the core  18  of the engine  10  from a front portion to a rear portion. The direction of the flow of gases through the core  18  of the engine  10  of  FIG.  1 B  can be better appreciated by considering that the gases flow through the core  18  in the same direction D as the one along which the engine  10  travels during flight for the engine. Stated differently, gases flow through the engine  10  of  FIG.  1 B  from a rear end towards a front end in the direction of the propeller  16 B. The direction of the flow of gases through the core  18  of the engines  10  of  FIGS.  1 A and  10    can be better appreciated by considering that the gases flow through the core  18  in a direction D1 that is opposite to the direction one along which the engines  10  travel during flight for the engines. Stated differently, gases flow through the engines  10  of  FIGS.  1 A and  10    from a front end towards a rear end in the direction of the exhaust outlet  15 . The engines  10  of  FIGS.  1 A to  10    may have one or multiple spools which perform compression to pressurize the air received through the air inlet  11 , and which extract energy from the combustion gases before they exit the core  18  via the exhaust outlet  15 . The spools and this engine architecture are described in greater detail in U.S. patent application Ser. No. 15/266,321 filed on Sep. 15, 2016, the entire contents of which are hereby incorporated by reference. 
     It will thus be appreciated that the expressions “forward” and “aft” used herein refer to the relative disposition of components of the engines  10 , in correspondence to the “forward” and “aft” directions of the engines  10  and aircraft including the engines  10  as defined with respect to the direction of travel. In  FIGS.  1 A and  10   , a component of the engines  10  that is “forward” of another component is arranged within the engine  10  such that it is located closer to the air inlet  11 . Similarly, a component of the engines  10  in  FIGS.  1 A and  10    that is “aft” of another component is arranged within the engines  10  such that it is further away from the air inlet  11 . In  FIG.  1 B , a component of the engine  10  that is “forward” of another component is arranged within the engine  10  such that it is located closer to the propeller  16 B. 
     Referring to  FIGS.  1 A to  10   , the air inlet  11  is the first point of entry of air into the core  18  of the engine  10 . The air inlet  11  has, or is defined by, an inlet duct  21  along which air flows as it drawn into the engine  10 . The inlet duct  21  may take different forms, as described in greater detail below. 
     Referring to  FIGS.  1 A and  1 B , the air inlet  11  is a radial air inlet  11  because, during operation of the engines  10 , air is drawn into the engine via the air inlet  11  along a substantially radial direction. The inlet duct  21  is defined by two annular walls  22 A,  22 B with sections that extend along substantially radial directions relative to the center axis  17 . Each wall  22 A,  22 B is shown as being an integral body. In an alternate embodiment, one or both of the walls  22 A,  22 B is made up of wall segments. Each annular wall  22 A,  22 B extends between a radially-outer portion  23 A and a radially-inner portion  23 B. The radially-inner portion  23 B is a portion of each wall  22 A,  22 B that is radially inward (i.e. closer to the center axis  17  of the engine  10 ) than the radially-outer portion  23 A. Each wall  22 A,  22 B therefore extends from an outer surface or portion of the engine  10  radially inwards toward the core  18 . The walls  22 A,  22 B in the depicted embodiment also have portions extending in an axial direction relative to the center axis  17 . The radially-inner portions  23 B of each wall  22 A,  22 B have trailing ends  24  which, in the frame of reference of the engine  10 , are defined by both axial and radial direction vectors. An air opening or inlet  25  is defined at the radially-outer portions  23 A of the walls  22 A,  22 B. The inlet  25  is circumferential because it spans a portion or all of the circumference of the inlet duct  21 . The inlet  25  extends through an outermost surface  26  of the engine  10 . The outermost surface  26  may be defined by an engine covering, such as a nacelle or casing. The inlet  25  may be provided with a screen, filter, or mesh to prevent the ingress of foreign objects into the engine  10 . The inlet duct  21  extends from the inlet  25  in a radially-inward direction to an outlet  24 A of the inlet duct  21  which is defined by the radially-inner portions  23 B of each wall  22 A,  22 B. The outlet  24 A is within the engines  10  and forms part of their cores  18 . 
     Referring to  FIGS.  1 A and  1 B , the walls  22 A,  22 B are axially spaced apart from one another. In  FIG.  1 A , the wall  22 B is aft of the wall  22 A in a direction along the center axis  17 . In  FIG.  1 B , the wall  22 B is forward of the wall  22 A in a direction along the center axis  17 . The axial offset between the annular walls  22 A,  22 B defines an inner volume of the inlet duct  21  through which air is conveyed toward the compressor section  12 . The spaced-apart walls  22 A,  22 B therefore define an annular air passage  27  between them. The air passage  27  is an annular volume that extends radially inwardly at the radially-outer portions  23 A and which has both axial and radial direction vectors at the radially-inner portion  23 B of the walls  22 A,  22 B. 
     Referring to  FIG.  10   , the air inlet  11  is an axial air inlet  11  because, during operation of the engine  10 , air is drawn into the engine via the air inlet  11  along a substantially axial direction. The inlet duct  21  is defined by an annular wall  22 A that extends along substantially an axial direction relative to the center axis  17 . The wall  22 A is shown as being an integral body. In an alternate embodiment, the wall  22 A is made up of wall segments. The annular wall  22 A extends between an axially-outer portion  23 A and an axially-inner portion  23 B. The axially-inner portion  23 B is a portion of the wall  22 A that is axially inward (i.e. closer to the compressor section  12  the engine  10 ) than the axially-outer portion  23 A. The wall  22 A therefore extends from an outer surface or portion of the engine  10  axially inwards toward the core  18 . An inlet  25  is defined at the axially-outer portion  23 A of the wall  22 A. The inlet  25  is circumferential because it spans a portion or all of the circumference of the inlet duct  21 . The wall  22 A defines an annular air passage  27 . The air passage  27  is an annular volume that extends axially inwardly at the axially-outer portions  23 A. The inlet duct  21  extends from the inlet  25  in a axially-inward direction to an outlet  24 A of the inlet duct  21  which is defined by the axially-inner portion  23 B of the wall  22 A. The outlet  24 A is within the engine  10  and forms part of the core  18 . 
     The air inlets  11  of the engines  10  of  FIGS.  1 A to  10    include structural supports, or struts  30 . The struts  30  may take different forms. 
     Referring to the radial air inlets  11  of  FIGS.  1 A and  1 B , multiple air inlet struts  30  are located within the inlet duct  21 . Each strut  30  is part of the fixed structure of the engine  10 . Each strut  30  is a stationary component that helps to provide structure to the air inlet  11 . The struts  30  are circumferentially spaced-apart from one another about the center axis  17  within the inlet duct  21 . Each strut  30  extends across the inlet duct  21  between the annular walls  22 A,  22 B and through the annular air passage  27 . Each strut  30  is attached to the annular walls  22 A,  22 B. In the depicted embodiment, each strut  30  is integral with the walls  22 A,  22 B. In an alternate embodiment, one or more of the struts  30  can be removably mounted to the walls  22 A,  22 B. Each of the struts  30  in the depicted embodiment is a radial air inlet strut  30  because it extends radially inwardly. Stated differently, each strut  30  has a radial span defined between a radially-outer edge which defines the leading edge  31 A of the strut  30  near the radially-outer portions  23 A of the walls  22 A,  22 B, and a radially-inner edge which defines the trailing edge  31 B near the radially-inner portions  23 B of the walls  22 A,  22 B. Some or all of the trailing edge  31 B is radially closer to the center axis  17  than the leading edge  31 A. The position of the edges  31 A,  31 B of the strut  30  relative to the engine  10  may vary, and what remains constant is that the trailing edge  31 B is downstream of the leading edge  31 A relative to the flow of air over the strut  30 . Referring to  FIGS.  1 A and  1 B , each strut  30  also has an axial span defined between the annular walls  22 A,  22 B of the inlet duct  21 . 
     Referring to  FIGS.  1 A and  1 B , one or more of the struts  30  is shaped like an airfoil. The airfoil shape of the strut  30  helps to guide the flow of air through the air inlet  11 . Each airfoil-shaped strut  30  includes the leading edge  31 A, and the trailing edge  31 B. The trailing edge  31 B is radially closer to the center axis  17  than the leading edge  31 A along some or all of its length. The strut  30  may be positioned radially inwardly of the inlet  25  and radially outwardly of the outlet  24 A. The strut  30  is positioned downstream of the inlet  25  and upstream of the outlet  24 A, relative to the direction of flow across the strut  30  from the leading edge  31 A to the trailing edge  31 B. In an embodiment, the strut  30  is positioned at or adjacent to the inlet  25 . The chord C of the strut  30  is therefore defined along a line extending between the leading and trailing edges  31 A,  31 B (see  FIGS.  2 A and  4 A ). The chord C therefore extends in a substantially radial direction. By “substantially radial”, it is understood that in the frame of reference of the engine  10 , the magnitude of the radial direction vector of the chord C may be much greater than the magnitude of the axial direction vector of the chord C. The chord C may have a camber or stagger angle. In alternate embodiments, one or more of the struts  30  do not have an airfoil shape. 
     Referring to  FIGS.  2 A and  2 B , one or more of the struts  30  has one or more internal strut passages  33 . Each strut passage  33  is a volume positioned within the body of the strut  30  that is sealed-off from the flow of air along the external surfaces of the strut  30 . The strut passage  33  allows for air to flow through the interior of the strut  30  in order to measure a static pressure at a location of the strut  30 , as explained in greater detail below. The strut passage  33  may be formed by drilling, etching, milling or any other operation for forming an internal volume within the material thickness of the strut  30 . Referring to  FIGS.  2 A and  2 B , the strut passage  33  extends to, through or is otherwise in fluid communication with, a pressure sensor  32 . The fluid communication between the pressure sensor  32  and the strut passage  33  allows the pressure sensor  32  to obtain a pressure reading from the air within the strut passage  33 . The pressure sensor  32  is fixedly mounted to the strut  30  or to any adjacent fixed structure using any suitable attachment technique. For example, and referring to  FIGS.  2 A and  2 B , the engine casing includes a boss  36  defining a groove for receiving the pressure sensor  32 . The pressure sensor  32  is attached to the engine inlet casing through the boss  36 , where the base of the groove of the boss  36  has an opening in fluid communication with the strut passage  33 . The boss  36  has an opening in fluid communication with the strut passage  33 . The internal strut passage  33  is thus in fluid communication with the pressure sensor  32  when it is mounted to the strut  30 . Referring to  FIGS.  2 A and  2 B , the strut passage  33  extends from a root of the strut  30  towards a tip of the strut  30 . In an alternate embodiment, the strut passage  33  is defined by a fluid line which extends along an external surface of the strut  30  to the pressure sensor  32 . In an alternate embodiment, the pressure sensor  32  is remotely mounted away from the strut  30  and engine casing. In such an alternate embodiment, a tube may extend from the boss  36  and be routed to a port of the pressure sensor  32  that is part of a control system. 
     The strut  30  has additional components which allow for a pressure reading of the air at locations on the strut  30  to be generated. Referring to  FIGS.  2 A and  2 B , one or more of the struts  30  has multiple static pressure measurement taps  38 . The static pressure measurement taps  38  allow the pressure sensor  32  to generate a reading of the static pressure at the static pressure measurement taps  38  (sometimes referred to herein simply as “taps  38 ”). In an embodiment, and referring to  FIGS.  2 A and  2 B , the taps  38  are used to obtain a reading of only the static pressure at the location of the taps  38 . The static pressure is the pressure applied by the air at the location of the taps  38  when the air has a substantially zero local velocity relative to the taps  38 . In an embodiment, the taps  38  exclude, prevent, or reduce the measurement of any dynamic pressure component of the air at the location of the taps  38 , where the dynamic pressure is the pressure applied by the air as a result of its motion relative to the taps  38 . In an embodiment, the taps  38  contribute to the measurement of a total or ram pressure component of the air at the location of the taps  38 , where the total pressure is the addition of static pressure and dynamic pressure at the taps  38 . 
     In an embodiment, the taps  38  capture the static pressure of the air at a location of the strut  30  where the dynamic pressure is approximately zero, such that the total pressure of the air at this location of the strut  30  is approximately equal to the local static pressure. Referring to  FIGS.  2 A and  2 B , such a location of the strut  30  may be its trailing edge  31 B. The taps  38  are small holes or openings that are spaced apart from each other along the span of the trailing edge  31 B, and which are formed in the radius or thickness of the trailing edge  31 B. The taps  38  are in fluid communication with the strut passage  33 . This allows the pressure sensor  32 , which is also in fluid communication with the strut passage  33  and thus in fluid communication with the taps  38 , to obtain a reading of the static pressure at the taps  38 . Referring to  FIG.  2 B , the static measurement taps  38  are positioned on the trailing edge  31 B of the top strut  30 , or the TDC strut  30 . In an embodiment, the TDC  30  is the only inlet strut  30  of the air inlet  11  that has taps  38  positioned on the trailing edge  31 B. Referring to  FIGS.  2 A  and  2 B, two taps  38  are shown at the trailing edge  31 B of the strut  30 , but more taps  38  may be used at the trailing edge  31 B. Using multiple taps  38  versus using a single tap  38  may help to provide a more reliable static pressure measurement, by helping to bring consistency to the static pressure measurement from an aerodynamics perspective by averaging the static pressure measurement over the multiple taps  38  which makes the measurement value less sensitive to the inlet flow variations at different flight conditions. Using multiple taps  38  versus using a single tap  38  may help to provide redundancy to the static pressure measurement, to allow for continuous measuring of the static pressure in the event that one or more the taps  38  ceases to function or becomes blocked. 
     By placing the taps  38  on the trailing edge  31 B of the strut  30 , the accuracy of the static pressure measurement may be improved by avoiding a local dynamic pressure contribution that would introduce error in the static pressure reading. The dynamic pressure contribution is expected to be suppressed or significantly minimized by locating the taps  38  at the trailing edge  31 B where the flow over the strut  30  is expected to be detached. Since the dynamic pressure contribution to the total pressure at the taps  38  may be minimal along the trailing edge  31 B, the static pressure measurement at the trailing edge  31 B may be used to determine the total pressure representative of the inlet pressure at the inlet of the compressor section  12 , sometimes referred to as the “P1” pressure in aircraft engines  10 . This P1 compressor inlet pressure may be more accurately captured using static pressure and may better resist the effects of inlet total pressure distortion that may reduce the accuracy of a system which directly measures the total pressure (i.e. such as a system using a pitot tube). 
     Positioning the static pressure measurement taps  38  on the trailing edge  31 B helps to orient the taps  38  away from the general direction of the flow of air over the strut  30 , and thereby decreases the likelihood that a dynamic pressure component will be added to the static pressure measurement even in scenarios where the flow or air remains attached. Positioning the static pressure measurement taps  38  on the trailing edge  31 B may also be beneficial from an icing perspective. Having two or more dedicated ports/taps  38  for predicting P1 provides redundancy, because multiple pressure measurement taps  38  on the trailing edge  31 B lowers the risk that ice will block all of the static pressure measurement taps  38  and prevent a proper reading of static pressure. Considering the redundancy and the low icing risk because the taps  38  are located on the trailing edge  31 B rather than along the leading edge  31 B which is more prone to icing, the taps  38  may not need to be located in the vicinity of a heat source. This may be beneficial for struts  30  used in radial air inlets  11 , such as the one shown in  FIG.  2 B , which may not have a dedicated de-icing capability and rely instead on an upstream inlet screen  35  to achieve de-icing. However, in an embodiment, the taps  38  are located in one or more struts  30  through which warm fluid (oil and/or air) passes through, and/or are in close proximity to a warmer bearing cavity. For example, and referring to  FIG.  2 B , the strut  30  shown is a top strut  30 , or a “top dead center” (TDC) strut  30 . The TDC strut  30  is located at the 12 o&#39;clock position when the aircraft engine  10  is mounted on the aircraft. The TDC strut  30  is located at the highest vertical position of all the struts  30  when the aircraft engine  10  is mounted on the aircraft. The TDC strut  30  benefits from some heating, such as by an oil pipe located close to the taps  38  which is used to transport used, warmed oil back to an oil tank. Alternatively, in the embodiment where the TDC strut  30  is present through an axial air inlet  11 , such as the one shown in  FIG.  10   , the taps  38  along the trailing edge  31 B may be in proximity to an internal bleed air passage extending through the TDC strut  30 . Such heat sources, through conduction, provide additional resiliency for de-icing. The risk of icing blocking the taps  38  may be further reduced by locating the taps  38  in a segment of the trailing edge  31 B where ice is less likely to accrete. The taps  38  may be provided with an active pressure tap protection system against icing, including a heating element. 
     The use of static pressure measurement taps  38  on the trailing edge  31 B of the inboard compressor inlet strut  30  to determine the total pressure (e.g. P1), as part of an engine control system, helps to accurately calculate the engine inlet, mass-averaged total pressure P1 over a range of engine conditions by measuring the static pressure at specific locations. The location of two or more static pressure taps  38  on the engine inlet case helps to provide means to measure the aerodynamically averaged pressure of those taps  38  and to determine the total pressure P1 from the static pressure measurements, and thus helps to provide a mass flow averaged total pressure that is representative of the pressure entering the compressor section  12 . 
     The use of static pressure measurement taps  38  on the trailing edge  31 B of the inboard compressor inlet strut  30  helps to improve P1 determination due to the accuracy and robustness of measuring static pressure at the trailing edge  31 B of the inlet strut  30 , when compared to other techniques used to estimate the inlet total pressure in flight. These other techniques, such as estimating P1 using ambient atmospheric pressure or an aircraft total pressure (pitot), do not capture the effect on the compressor inlet pressure of various operational or installation effects such as inlet pressure losses due to icing, variations in angle of attack, inlet by-pass flow, inertial particle separators, inlet barrier filters, and/or left/right/center installation of the engine  10  on the aircraft. The use of static pressure measurement taps  38  on the trailing edge  31 B of the inboard compressor inlet strut  30  determines the static pressure downstream of many or all of these operational or installation effects, and consequently helps to provide a more accurate input of P1 to the engine control system by measuring at a location very close to the inlet of the compressor section  12 , or at a location which can be correlated with high accuracy to the inlet of the compressor section  12 . This may help to improve the engine performance, operability and controls system. Examples of control system logics and algorithm which may be improved by such a more accurate input include, but are not limited to: variable guide vane controls, handling bleed valve controls, engine limiting loops, turbine temperature algorithm, power assurance checks, and power settings. 
     The use of static pressure measurement taps  38  on the trailing edge  31 B of the inboard compressor inlet strut  30  allows for determining the total pressure P1 in compressor inlet flows that are distorted, such as those through radial air inlets  11 . The total air inlet pressure P1 may be determined indirectly, via the use of the measured static pressure at the taps  38 , thereby allowing an accurate measure of static pressure to be converted or correlated reliably to a total pressure value, such as P1. 
     Referring to  FIG.  2 B , the trailing edge  31 B of the strut  30  has multiple edge contours  39 . Each edge contour  39  is a localised feature on the trailing edge  31 B, and is positioned at one of the taps  38 . By “positioned”, it is understood that each edge contour  39  is associated with one of the taps  38 , and surrounds or encircles the tap  38 . The edge contours  39  help the flow over the strut  30  to separate at the trailing edge  31 B, thereby contributing to allowing the taps  38  to adequately capture the static pressure at the location of the taps  38 . Each edge contour  39  helps to achieve this function by forming, or being defined by, a contour edge wall  39 W that is recessed from a remainder of a surface  31 BS of the trailing edge  31 B. The recession of the contour edge wall  39 W may take different forms. For example, and referring to  FIGS.  2 B and  3   , each contour edge wall  39 W is a curved wall that extends from the adjacent, unaltered surface  31 BS of the trailing edge  31 B into the body of the strut  30  (i.e. towards the leading edge  31 A). For example, and referring to  FIGS.  2 B and  3   , each contour edge wall  39 W is a hemispherical cut-out of material from the surface  31 BS of the trailing edge  31 B. Referring to  FIGS.  2 B and  3   , the taps  38  are openings defined in the contour edge walls  39 W, and the taps  38  are also recessed from the remaining surface  31 BS of the trailing edge  31 B. Thus, in the configuration of  FIGS.  2 B and  3   , each contour edge wall  39 W is a divot or cut-out of the trailing edge  31 B that is localized around the opening in the trailing edge  31 B that forms the tap  38 . Referring to  FIGS.  2 B and  3   , the curved contour edge wall  39 W spaces the tap  38  from the remaining surface  31 BS of the trailing edge  31 B, such that a depth distance greater than zero is defined between the remaining surface  31 BS and the tap  38 . 
     Referring to  FIG.  4 B , another configuration of the contour edge wall  39 W of the edge contours  39  is shown. Each contour edge wall  39 W includes two wall sections that extend from the adjacent, unaltered surface  31 BS of the trailing edge  31 B into the body of the strut  30  (i.e. towards the leading edge  31 A). Each contour edge wall  39 W is a cut-out of material from the surface  31 BS of the trailing edge  31 B. Each contour edge wall  39 W also includes a planar wall section extending between the two wall sections. The planar wall section is recessed from the remaining surface  31 BS by the two wall sections. The taps  38  are openings defined in the planar wall section of the contour edge walls  39 W, and the taps  38  are also recessed from the remaining surface  31 BS of the trailing edge  31 B. Thus, in the configuration of  FIG.  4 B , each contour edge wall  39 W is a divot or cut-out of the trailing edge  31 B that is localized around the opening in the trailing edge  31 B that forms the tap  38 . The two wall sections of the contour edge wall  39 W space the tap  38  from the remaining surface  31 BS of the trailing edge  31 B, such that a depth distance greater than zero is defined between the remaining surface  31 BS and the planar wall section/tap  38 . 
     Irrespective of its configuration, the edge contour  39  may be referred to as a “drip edge” because its recessed shape may assist in preventing water from running back into the tap  38  and subsequently freezing to thereby block the tap  38 . In some situations, the edge contour  39  helps to “sharpen” the trailing edge  31 B at the location of the edge contour  39 , which may result in water have a greater tendency to be shed from the “sharper” trailing edge  31 B and not enter the tap  38 , compared to a trailing edge  31 B that is more rounded. From an aerodynamics perspective, such a drip edge  39  may also improve the accuracy of the static pressure measurement at the tap  38  by forcing the flow to separate at the trailing edge  31 B in the vicinity of the taps  38 . For some flight conditions, the inlet flow entering the air passage  27  at a high incidence with respect to the strut  30  may modify the expected location of the flow separation at the trailing edge  31 B. Thus, a significant flow angle of attack with respect to the strut  30  may result in a separation delay at the trailing edge  31 B. In this situation, the static pressure measurement may be skewed by a dynamic pressure contribution. The edge contour  39  may counter this effect by forcing the flow to separate at the trailing edge  31 B in the vicinity of the taps  38 . Furthermore, there may be an inherent transient behaviour of the flow that could lead to undesired measurement fluctuations and increase the uncertainty about the measurement quality. The edge contour  39  may help prevent such uncertainty by forcing the flow detachment at the trailing edge  31 B. 
     As described above, the strut passage  33  is in fluid communication with the taps  38  on the trailing edge  31 B. This fluid communication arrangement may take different forms. For example, and referring to  FIG.  2 B , the strut  30  includes multiple measurement tap passages  37  that extend within the strut  30  and which are not exposed to the flow of air over the external surfaces of the strut  30 . The measurement tap passages  37  may be formed by any suitable technique that removes material from within the strut  30 , such as drilling, milling or etching. Each measurement tap passage  37  extends within the strut  30  in a direction toward the trailing edge  31 B and one or more of the taps  38 , thereby fluid linking those taps  38  to the strut passage  33  and to the pressure sensor  32 . Referring to  FIG.  2 B , the diameter of the measurement tap passages  37  is less than the diameter of the strut passage  33 . The measurement tap passages  37  thus establish fluid communication between the taps  38  on the trailing edge  31 B and the pressure sensor  32 , via the strut passage  33 . 
     The fluid communication between the strut passage  33  and the taps  38  takes another form in an alternate embodiment. In this alternate embodiment, the strut passage  33  is composed of two or more strut passages  33 . Each strut passage  33  is in fluid communication with one of the taps  38  via a dedicated measurement tap passage  37 . Each strut passage  33  is fluidly connected to its own pressure sensor  32 . Thus, in this alternate embodiment, the taps  38  are independent of one another, and each of the taps  38  has their own pressure sensor  32 . 
     The internal passages  33 ,  37  of the strut  30  may be shaped or oriented to facilitate fluid drainage within the strut  30 , or to reduce or prevent fluid accumulation within the strut  30 . For example, and referring to  FIG.  2 B , the internal strut passage  33  and the measurement tap passages  37  slope toward the center axis  17  of the aircraft engine  10 . In the configuration of the strut  30  shown in  FIG.  2 B , each of the measurement tap passages  37  has an orientation defined by a predominant directional vector that is radial relative to the center axis  17 , such that any water which may be present within the measurement tap passages  37  would flow due to gravity toward the trailing edge  31 B and out of the tap  38 . In the configuration of the strut  30  shown in  FIG.  2 B , drainage of the strut passage  33  is achieved by positioning the strut passage  33  so that it is radially outwardly of the location where it intersects each of the measurement tap passages  37 . In the configuration of the strut  30  shown in  FIG.  2 B , the strut passage has  33  a lowermost point  33 LP that is the vertically lowest portion of the strut passage  33  when the engine  10  is mounted on the aircraft. The lowermost point  33 LP is the portion of strut passage  33  that is radially closest to the center axis  17 . The lowermost point  33 LP is positioned at the highest point (i.e. radially-outermost) of the lowermost measurement tap passage  37 L. Referring to  FIG.  2 B , the end of the strut passage  33  opposite to the end engaged with the boss  36  is plugged or sealed. 
     In the configuration of the strut  30  shown in  FIGS.  4 A and  4 B , each of the measurement tap passages  37  has an orientation defined by a predominant directional vector that is radial relative to the center axis  17 , such that any water which may be present within the measurement tap passages  37  would flow due to gravity toward the trailing edge  31 B and out of the tap  38 . In the configuration of the strut  30  shown in  FIGS.  4 A and  4 B , drainage of the strut passage  33  is achieved by positioning the strut passage  33  so that it is radially outwardly of the location where it intersects each of the measurement tap passage  37 . In the configuration of the strut  30  shown in  FIGS.  4 A and  4 B , the strut passage has  33  a lowermost point  33 LP that is the vertically lowest portion of the strut passage  33  when the engine  10  is mounted on the aircraft. The lowermost point  33 LP is the portion of strut passage  33  that is radially closest to the center axis  17 . The lowermost point  33 LP allows any moisture within the strut passage  33  to accumulate at the lowermost point  33 LP, and is positioned adjacent to the highest point (i.e. radially-outermost) of the lowermost measurement tap passage  37 L such that water can flow from the lowermost point  33 LP into, and then out of, the lowermost measurement tap passage  37 L. 
     These configurations of the strut passage  33  and of the measurement tap passages  37  facilitate gravity drainage of moisture that may accumulate within the strut passage  33 , which may prevent accurate static pressure measurements at the taps  38  if the moisture is left undrained, or if the moisture freezes within the internal passages  33 ,  37 . The internal passages  33 ,  37  of the strut  30  may thus have any suitable angle with respect to the attitude of the aircraft on the ground to facilitate drainage. 
     Referring to  FIGS.  2 A and  2 B , the strut  30  has a first static pressure measurement tap  38 A in fluid communication with the lowermost measurement tap passage  37 L, and a second static pressure measurement tap  38 B in fluid communication with the other measurement tap passage  37 . The first tap  38 A is positioned adjacent to one of a root and a tip of the strut  30  and trailing edge  31 B, such as adjacent to the root in  FIGS.  2 A and  2 B . The second tap  38 B is positioned adjacent to the other of the root and the tip, such as adjacent to the tip of the trailing edge  31 B in  FIGS.  2 A and  2 B . The trailing edge  31 B has a span S measured between the root and the tip. The distance separating the first and second taps  38 A,  38 B along the span S of the trailing edge  31 B is greater than zero. The distance separating the first and second taps  38 A,  38 B along the span S of the trailing edge  31 B is not less than 40% of the span S. Keeping such a minimum spanwise distance between the taps  38 A,  38 B may minimize their sensitivity and improve the accuracy of averaging the measurement of static pressure at their locations. Alternatively, the distance separating the taps  38 A,  38 B may be less than 40% of the span S depending on the aerodynamics at the trailing edge  31 B for various flight conditions, in order to have a consistently accurate P1 calculation through a wide range of engine powers and flight conditions. Referring to  FIG.  2 B , a length L of each measurement tap passage  37  is defined between a first end  37 A at the strut passage  30  and a second end  37 B at the trailing edge  31 B of the strut  30 . A ratio of the length L over a diameter of each of the openings which form the taps  38  is greater than three. Such a ratio may help to reduce any measurement error in the static pressure. From an aerodynamics perspective, such a minimum L/D ratio may help to minimize the flow turbulence at the taps  38 . The radius of each tap  38  may be sized with respect to the radius of the trailing edge  31 B. For example, radius of each tap  38  may be between ⅕ to ½ of the radius of the trailing edge  31 B. Having a bigger hole diameter for the tap  38  may make the tap  38  more tolerant to the ingestion of a few water droplets because it mitigates the risk of having some of these water droplets trapped in the measurement tap passage  37  and potentially freeze. Selecting the diameter of the taps  38  may be a trade-off between the aerodynamics and icing requirements on one side, and the mechanical and manufacturing constraints on the other side. 
     Although the strut  30 , the strut passage  33 , the edge contour  39 , the measurement tap passages  37 , and the static pressure measurement taps  38  are described in relation to an inlet strut  30  in a radial air inlet  11 , the description and associated advantages of these features shown in  FIGS.  2 A and  2 B  applies mutatis mutandis to a strut  30  present in an axial air inlet  11 , such as the one shown in  FIG.  10   . Similarly, the description and associated advantages of the strut  30 , the strut passage  33 , the edge contour  39 , the measurement tap passages  37 , and the static pressure measurement taps  38  shown in  FIGS.  2 A and  2 B  applies mutatis mutandis to the strut  30  present in the radial air inlet  11  of  FIGS.  1 B,  4 A and  4 B . 
     Referring to  FIGS.  2 A and  2 B , there is disclosed a method of obtaining static pressure in an inlet of a compressor of an aircraft engine. The method includes obtaining the static pressure at spaced-apart locations on a trailing edge of a strut extending across the inlet. 
     The embodiments described in this document provide non-limiting examples of possible implementations of the present technology. Upon review of the present disclosure, a person of ordinary skill in the art will recognize that changes may be made to the embodiments described herein without departing from the scope of the present technology. Yet further modifications could be implemented by a person of ordinary skill in the art in view of the present disclosure, which modifications would be within the scope of the present technology.