Patent Description:
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.

<CIT> discloses a gas turbine engine comprising an air inlet duct and a strut. <CIT> discloses a pitot-static tube with static orifices on a plate upstream of a strut, in which at least one orifice for tapping static pressure is arranged on an exit section of said pitot-static tube.

According to a first aspect of the invention there is an aircraft engine as claimed in claim <NUM>. Embodiments of this aspect of the invention are as claimed in the dependent claims.

<FIG> illustrate different gas turbine engines <NUM> of a type preferably provided for use in subsonic flight. Each of the gas turbine engines <NUM> generally comprises in serial flow communication an air inlet <NUM>, a compressor section <NUM> for pressurizing the air from the air inlet <NUM>, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, a turbine section <NUM> for extracting energy from the combustion gases, and an exhaust outlet <NUM> through which the combustion gases exit the gas turbine engine <NUM>. The gas turbine engine <NUM> have a longitudinal center axis <NUM> about which components rotate. In the gas turbine engines <NUM> shown in <FIG> and <FIG>, the air inlet <NUM> is positioned forward of the compressor section <NUM>, whereas in the gas turbine engine <NUM> shown in <FIG>, the air inlet <NUM> is positioned aft of the compressor section <NUM>. The gas turbine engine <NUM> of <FIG> includes a driven gear train 16A mounted at a front end of the gas turbine engine <NUM>, and is an example of a "turboshaft" gas turbine engine <NUM>. The gas turbine engine <NUM> of <FIG> includes a propeller 16B which provides thrust for flight and taxiing, and is an example of a "turboprop" gas turbine engine <NUM>. The gas turbine engine <NUM> of <FIG> includes a fan 16C which provides thrust for flight, and is an example of a "turbofan" gas turbine engine <NUM>.

The gas turbine engines <NUM> (sometimes referred to herein simply as "engines <NUM>") have a central core <NUM> through which gases flow and which includes some of the turbomachinery of the engine <NUM>. The engine <NUM> of <FIG> is a "reverse-flow" engine <NUM> because gases flow through the core <NUM> from the air inlet <NUM> at a rear portion, to the exhaust outlet <NUM> at a front portion. This is in contrast to "through-flow" gas turbine engines <NUM>, such as those shown in <FIG> and <FIG>, in which gases flow through the core <NUM> of the engine <NUM> from a front portion to a rear portion. The direction of the flow of gases through the core <NUM> of the engine <NUM> of <FIG> can be better appreciated by considering that the gases flow through the core <NUM> in the same direction D as the one along which the engine <NUM> travels during flight for the engine. Stated differently, gases flow through the engine <NUM> of <FIG> from a rear end towards a front end in the direction of the propeller 16B. The direction of the flow of gases through the core <NUM> of the engines <NUM> of <FIG> and <FIG> can be better appreciated by considering that the gases flow through the core <NUM> in a direction D1 that is opposite to the direction one along which the engines <NUM> travel during flight for the engines. Stated differently, gases flow through the engines <NUM> of <FIG> and <FIG> from a front end towards a rear end in the direction of the exhaust outlet <NUM>. The engines <NUM> of <FIG> may have one or multiple spools which perform compression to pressurize the air received through the air inlet <NUM>, and which extract energy from the combustion gases before they exit the core <NUM> via the exhaust outlet <NUM>. The spools and this engine architecture are described in greater detail in <CIT>.

It will thus be appreciated that the expressions "forward" and "aft" used herein refer to the relative disposition of components of the engines <NUM>, in correspondence to the "forward" and "aft" directions of the engines <NUM> and aircraft including the engines <NUM> as defined with respect to the direction of travel. In <FIG> and <FIG>, a component of the engines <NUM> that is "forward" of another component is arranged within the engine <NUM> such that it is located closer to the air inlet <NUM>. Similarly, a component of the engines <NUM> in <FIG> and <FIG> that is "aft" of another component is arranged within the engines <NUM> such that it is further away from the air inlet <NUM>. In <FIG>, a component of the engine <NUM> that is "forward" of another component is arranged within the engine <NUM> such that it is located closer to the propeller 16B.

Referring to <FIG>, the air inlet <NUM> is the first point of entry of air into the core <NUM> of the engine <NUM>. The air inlet <NUM> has, or is defined by, an inlet duct <NUM> along which air flows as it drawn into the engine <NUM>. The inlet duct <NUM> may take different forms, as described in greater detail below.

Referring to <FIG>, the air inlet <NUM> is a radial air inlet <NUM> because, during operation of the engines <NUM>, air is drawn into the engine via the air inlet <NUM> along a substantially radial direction. The inlet duct <NUM> is defined by two annular walls 22A,22B with sections that extend along substantially radial directions relative to the center axis <NUM>. Each wall 22A,22B is shown as being an integral body. In an alternate embodiment, one or both of the walls 22A,22B is made up of wall segments. Each annular wall 22A,22B extends between a radially-outer portion 23A and a radially-inner portion 23B. The radially-inner portion 23B is a portion of each wall 22A,22B that is radially inward (i.e. closer to the center axis <NUM> of the engine <NUM>) than the radially-outer portion 23A. Each wall 22A,22B therefore extends from an outer surface or portion of the engine <NUM> radially inwards toward the core <NUM>. The walls 22A,22B in the depicted embodiment also have portions extending in an axial direction relative to the center axis <NUM>. The radially-inner portions 23B of each wall 22A,22B have trailing ends <NUM> which, in the frame of reference of the engine <NUM>, are defined by both axial and radial direction vectors. An air opening or inlet <NUM> is defined at the radially-outer portions 23A of the walls 22A,22B. The inlet <NUM> is circumferential because it spans a portion or all of the circumference of the inlet duct <NUM>. The inlet <NUM> extends through an outermost surface <NUM> of the engine <NUM>. The outermost surface <NUM> may be defined by an engine covering, such as a nacelle or casing. The inlet <NUM> may be provided with a screen, filter, or mesh to prevent the ingress of foreign objects into the engine <NUM>. The inlet duct <NUM> extends from the inlet <NUM> in a radially-inward direction to an outlet 24A of the inlet duct <NUM> which is defined by the radially-inner portions 23B of each wall 22A,22B. The outlet 24A is within the engines <NUM> and forms part of their cores <NUM>.

Referring to <FIG>, the walls 22A,22B are axially spaced apart from one another. In <FIG>, the wall 22B is aft of the wall 22A in a direction along the center axis <NUM>. In <FIG>, the wall 22B is forward of the wall 22A in a direction along the center axis <NUM>. The axial offset between the annular walls 22A,22B defines an inner volume of the inlet duct <NUM> through which air is conveyed toward the compressor section <NUM>. The spaced-apart walls 22A,22B therefore define an annular air passage <NUM> between them. The air passage <NUM> is an annular volume that extends radially inwardly at the radially-outer portions 23A and which has both axial and radial direction vectors at the radially-inner portion 23B of the walls 22A,22B.

Referring to <FIG>, the air inlet <NUM> is an axial air inlet <NUM> because, during operation of the engine <NUM>, air is drawn into the engine via the air inlet <NUM> along a substantially axial direction. The inlet duct <NUM> is defined by an annular wall 22A that extends along substantially an axial direction relative to the center axis <NUM>. The wall 22A is shown as being an integral body. In an alternate embodiment, the wall 22A is made up of wall segments. The annular wall 22A extends between an axially-outer portion 23A and an axially-inner portion 23B. The axially-inner portion 23B is a portion of the wall 22A that is axially inward (i.e. closer to the compressor section <NUM> the engine <NUM>) than the axially-outer portion 23A. The wall 22A therefore extends from an outer surface or portion of the engine <NUM> axially inwards toward the core <NUM>. An inlet <NUM> is defined at the axially-outer portion 23A of the wall 22A. The inlet <NUM> is circumferential because it spans a portion or all of the circumference of the inlet duct <NUM>. The wall 22A defines an annular air passage <NUM>. The air passage <NUM> is an annular volume that extends axially inwardly at the axially-outer portions 23A. The inlet duct <NUM> extends from the inlet <NUM> in a axially-inward direction to an outlet 24A of the inlet duct <NUM> which is defined by the axially-inner portion 23B of the wall 22A. The outlet 24A is within the engine <NUM> and forms part of the core <NUM>.

The air inlets <NUM> of the engines <NUM> of <FIG> include structural supports, or struts <NUM>. The struts <NUM> may take different forms.

Referring to the radial air inlets <NUM> of <FIG>, multiple air inlet struts <NUM> are located within the inlet duct <NUM>. Each strut <NUM> is part of the fixed structure of the engine <NUM>. Each strut <NUM> is a stationary component that helps to provide structure to the air inlet <NUM>. The struts <NUM> are circumferentially spaced-apart from one another about the center axis <NUM> within the inlet duct <NUM>. Each strut <NUM> extends across the inlet duct <NUM> between the annular walls 22A,22B and through the annular air passage <NUM>. Each strut <NUM> is attached to the annular walls 22A,22B. In the depicted embodiment, each strut <NUM> is integral with the walls 22A,22B. In an alternate embodiment, one or more of the struts <NUM> can be removably mounted to the walls 22A,22B. Each of the struts <NUM> in the depicted embodiment is a radial air inlet strut <NUM> because it extends radially inwardly. Stated differently, each strut <NUM> has a radial span defined between a radially-outer edge which defines the leading edge 31A of the strut <NUM> near the radially-outer portions 23A of the walls 22A,22B, and a radially-inner edge which defines the trailing edge 31B near the radially-inner portions 23B of the walls 22A,22B. Some or all of the trailing edge 31B is radially closer to the center axis <NUM> than the leading edge 31A. The position of the edges 31A,31B of the strut <NUM> relative to the engine <NUM> may vary, and what remains constant is that the trailing edge 31B is downstream of the leading edge 31A relative to the flow of air over the strut <NUM>. Referring to <FIG>, each strut <NUM> also has an axial span defined between the annular walls 22A,22B of the inlet duct <NUM>.

Referring to <FIG>, one or more of the struts <NUM> is shaped like an airfoil. The airfoil shape of the strut <NUM> helps to guide the flow of air through the air inlet <NUM>. Each airfoil-shaped strut <NUM> includes the leading edge 31A, and the trailing edge 31B. The trailing edge 31B is radially closer to the center axis <NUM> than the leading edge 31A along some or all of its length. The strut <NUM> may be positioned radially inwardly of the inlet <NUM> and radially outwardly of the outlet 24A. The strut <NUM> is positioned downstream of the inlet <NUM> and upstream of the outlet 24A, relative to the direction of flow across the strut <NUM> from the leading edge 31A to the trailing edge 31B. In an embodiment, the strut <NUM> is positioned at or adjacent to the inlet <NUM>. The chord C of the strut <NUM> is therefore defined along a line extending between the leading and trailing edges 31A,31B (see <FIG> and <FIG>). 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 <NUM>, 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 <NUM> do not have an airfoil shape.

Referring to <FIG>, one or more of the struts <NUM> has one or more internal strut passages <NUM>. Each strut passage <NUM> is a volume positioned within the body of the strut <NUM> that is sealed-off from the flow of air along the external surfaces of the strut <NUM>. The strut passage <NUM> allows for air to flow through the interior of the strut <NUM> in order to measure a static pressure at a location of the strut <NUM>, as explained in greater detail below. The strut passage <NUM> may be formed by drilling, etching, milling or any other operation for forming an internal volume within the material thickness of the strut <NUM>. Referring to <FIG>, the strut passage <NUM> extends to, through or is otherwise in fluid communication with, a pressure sensor <NUM>. The fluid communication between the pressure sensor <NUM> and the strut passage <NUM> allows the pressure sensor <NUM> to obtain a pressure reading from the air within the strut passage <NUM>. The pressure sensor <NUM> is fixedly mounted to the strut <NUM> or to any adjacent fixed structure using any suitable attachment technique. For example, and referring to <FIG>, the engine casing includes a boss <NUM> defining a groove for receiving the pressure sensor <NUM>. The pressure sensor <NUM> is attached to the engine inlet casing through the boss <NUM>, where the base of the groove of the boss <NUM> has an opening in fluid communication with the strut passage <NUM>. The boss <NUM> has an opening in fluid communication with the strut passage <NUM>. The internal strut passage <NUM> is thus in fluid communication with the pressure sensor <NUM> when it is mounted to the strut <NUM>. Referring to <FIG>, the strut passage <NUM> extends from a root of the strut <NUM> towards a tip of the strut <NUM>. In an alternate embodiment, the strut passage <NUM> is defined by a fluid line which extends along an external surface of the strut <NUM> to the pressure sensor <NUM>. In an alternate embodiment, the pressure sensor <NUM> is remotely mounted away from the strut <NUM> and engine casing. In such an alternate embodiment, a tube may extend from the boss <NUM> and be routed to a port of the pressure sensor <NUM> that is part of a control system.

The strut <NUM> has additional components which allow for a pressure reading of the air at locations on the strut <NUM> to be generated. Referring to <FIG>, one or more of the struts <NUM> has multiple static pressure measurement taps <NUM>. The static pressure measurement taps <NUM> allow the pressure sensor <NUM> to generate a reading of the static pressure at the static pressure measurement taps <NUM> (sometimes referred to herein simply as "taps <NUM>"). In an embodiment, and referring to <FIG>, the taps <NUM> are used to obtain a reading of only the static pressure at the location of the taps <NUM>. The static pressure is the pressure applied by the air at the location of the taps <NUM> when the air has a substantially zero local velocity relative to the taps <NUM>. In an embodiment, the taps <NUM> exclude, prevent, or reduce the measurement of any dynamic pressure component of the air at the location of the taps <NUM>, where the dynamic pressure is the pressure applied by the air as a result of its motion relative to the taps <NUM>. In an embodiment, the taps <NUM> contribute to the measurement of a total or ram pressure component of the air at the location of the taps <NUM>, where the total pressure is the addition of static pressure and dynamic pressure at the taps <NUM>.

In an embodiment, the taps <NUM> capture the static pressure of the air at a location of the strut <NUM> where the dynamic pressure is approximately zero, such that the total pressure of the air at this location of the strut <NUM> is approximately equal to the local static pressure. Referring to <FIG>, such a location of the strut <NUM> may be its trailing edge 31B. The taps <NUM> are small holes or openings that are spaced apart from each other along the span of the trailing edge 31B, and which are formed in the radius or thickness of the trailing edge 31B. The taps <NUM> are in fluid communication with the strut passage <NUM>. This allows the pressure sensor <NUM>, which is also in fluid communication with the strut passage <NUM> and thus in fluid communication with the taps <NUM>, to obtain a reading of the static pressure at the taps <NUM>. Referring to <FIG>, the static measurement taps <NUM> are positioned on the trailing edge 31B of the top strut <NUM>, or the TDC strut <NUM>. In an embodiment, the TDC <NUM> is the only inlet strut <NUM> of the air inlet <NUM> that has taps <NUM> positioned on the trailing edge 31B. Referring to <FIG>, two taps <NUM> are shown at the trailing edge 31B of the strut <NUM>, but more taps <NUM> may be used at the trailing edge 31B. Using multiple taps <NUM> versus using a single tap <NUM> 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 <NUM> which makes the measurement value less sensitive to the inlet flow variations at different flight conditions. Using multiple taps <NUM> versus using a single tap <NUM> 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 <NUM> ceases to function or becomes blocked.

By placing the taps <NUM> on the trailing edge 31B of the strut <NUM>, 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 <NUM> at the trailing edge 31B where the flow over the strut <NUM> is expected to be detached. Since the dynamic pressure contribution to the total pressure at the taps <NUM> may be minimal along the trailing edge 31B, the static pressure measurement at the trailing edge 31B may be used to determine the total pressure representative of the inlet pressure at the inlet of the compressor section <NUM>, sometimes referred to as the "P1" pressure in aircraft engines <NUM>. 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 <NUM> on the trailing edge 31B helps to orient the taps <NUM> away from the general direction of the flow of air over the strut <NUM>, 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 <NUM> on the trailing edge 31B may also be beneficial from an icing perspective. Having two or more dedicated ports/taps <NUM> for predicting P1 provides redundancy, because multiple pressure measurement taps <NUM> on the trailing edge 31B lowers the risk that ice will block all of the static pressure measurement taps <NUM> and prevent a proper reading of static pressure. Considering the redundancy and the low icing risk because the taps <NUM> are located on the trailing edge 31B rather than along the leading edge 31A which is more prone to icing, the taps <NUM> may not need to be located in the vicinity of a heat source. This may be beneficial for struts <NUM> used in radial air inlets <NUM>, such as the one shown in <FIG>, which may not have a dedicated de-icing capability and rely instead on an upstream inlet screen <NUM> to achieve de-icing. However, in an embodiment, the taps <NUM> are located in one or more struts <NUM> 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>, the strut <NUM> shown is a top strut <NUM>, or a "top dead center" (TDC) strut <NUM>. The TDC strut <NUM> is located at the <NUM> o'clock position when the aircraft engine <NUM> is mounted on the aircraft. The TDC strut <NUM> is located at the highest vertical position of all the struts <NUM> when the aircraft engine <NUM> is mounted on the aircraft. The TDC strut <NUM> benefits from some heating, such as by an oil pipe located close to the taps <NUM> which is used to transport used, warmed oil back to an oil tank. Alternatively, in the embodiment where the TDC strut <NUM> is present through an axial air inlet <NUM>, such as the one shown in <FIG>, the taps <NUM> along the trailing edge 31B may be in proximity to an internal bleed air passage extending through the TDC strut <NUM>. Such heat sources, through conduction, provide additional resiliency for de-icing. The risk of icing blocking the taps <NUM> may be further reduced by locating the taps <NUM> in a segment of the trailing edge 31B where ice is less likely to accrete. The taps <NUM> may be provided with an active pressure tap protection system against icing, including a heating element.

The use of static pressure measurement taps <NUM> on the trailing edge 31B of the inboard compressor inlet strut <NUM> 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 <NUM> on the engine inlet case helps to provide means to measure the aerodynamically averaged pressure of those taps <NUM> 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 <NUM>.

The use of static pressure measurement taps <NUM> on the trailing edge 31B of the inboard compressor inlet strut <NUM> helps to improve P1 determination due to the accuracy and robustness of measuring static pressure at the trailing edge 31B of the inlet strut <NUM>, 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 <NUM> on the aircraft. The use of static pressure measurement taps <NUM> on the trailing edge 31B of the inboard compressor inlet strut <NUM> 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 <NUM>, or at a location which can be correlated with high accuracy to the inlet of the compressor section <NUM>. 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 <NUM> on the trailing edge 31B of the inboard compressor inlet strut <NUM> allows for determining the total pressure P1 in compressor inlet flows that are distorted, such as those through radial air inlets <NUM>. The total air inlet pressure P1 may be determined indirectly, via the use of the measured static pressure at the taps <NUM>, 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>, the trailing edge 31B of the strut <NUM> has multiple edge contours <NUM>. Each edge contour <NUM> is a localised feature on the trailing edge 31B, and is positioned at one of the taps <NUM>. By "positioned", it is understood that each edge contour <NUM> is associated with one of the taps <NUM>, and surrounds or encircles the tap <NUM>. The edge contours <NUM> help the flow over the strut <NUM> to separate at the trailing edge 31B, thereby contributing to allowing the taps <NUM> to adequately capture the static pressure at the location of the taps <NUM>. Each edge contour <NUM> helps to achieve this function by forming, or being defined by, a contour edge wall 39W that is recessed from a remainder of a surface 31BS of the trailing edge 31B. The recession of the contour edge wall 39W may take different forms. For example, and referring to <FIG> and <FIG>, each contour edge wall 39W is a curved wall that extends from the adjacent, unaltered surface 31BS of the trailing edge 31B into the body of the strut <NUM> (i.e. towards the leading edge 31A). For example, and referring to <FIG> and <FIG>, each contour edge wall 39W is a hemispherical cut-out of material from the surface 31BS of the trailing edge 31B. Referring to <FIG> and <FIG>, the taps <NUM> are openings defined in the contour edge walls 39W, and the taps <NUM> are also recessed from the remaining surface 31BS of the trailing edge 31B. Thus, in the configuration of <FIG> and <FIG>, each contour edge wall 39W is a divot or cut-out of the trailing edge 31B that is localized around the opening in the trailing edge 31B that forms the tap <NUM>. Referring to <FIG> and <FIG>, the curved contour edge wall 39W spaces the tap <NUM> from the remaining surface 31BS of the trailing edge 31B, such that a depth distance greater than zero is defined between the remaining surface 31BS and the tap <NUM>.

Referring to <FIG>, another configuration of the contour edge wall 39W of the edge contours <NUM> is shown. Each contour edge wall 39W includes two wall sections that extend from the adjacent, unaltered surface 31BS of the trailing edge 31B into the body of the strut <NUM> (i.e. towards the leading edge 31A). Each contour edge wall 39W is a cut-out of material from the surface 31BS of the trailing edge 31B. Each contour edge wall 39W also includes a planar wall section extending between the two wall sections. The planar wall section is recessed from the remaining surface 31BS by the two wall sections. The taps <NUM> are openings defined in the planar wall section of the contour edge walls 39W, and the taps <NUM> are also recessed from the remaining surface 31BS of the trailing edge 31B. Thus, in the configuration of <FIG>, each contour edge wall 39W is a divot or cut-out of the trailing edge 31B that is localized around the opening in the trailing edge 31B that forms the tap <NUM>. The two wall sections of the contour edge wall 39W space the tap <NUM> from the remaining surface 31BS of the trailing edge 31B, such that a depth distance greater than zero is defined between the remaining surface 31BS and the planar wall section/tap <NUM>.

Irrespective of its configuration, the edge contour <NUM> may be referred to as a "drip edge" because its recessed shape may assist in preventing water from running back into the tap <NUM> and subsequently freezing to thereby block the tap <NUM>. In some situations, the edge contour <NUM> helps to "sharpen" the trailing edge 31B at the location of the edge contour <NUM>, which may result in water have a greater tendency to be shed from the "sharper" trailing edge 31B and not enter the tap <NUM>, compared to a trailing edge 31B that is more rounded. From an aerodynamics perspective, such a drip edge <NUM> may also improve the accuracy of the static pressure measurement at the tap <NUM> by forcing the flow to separate at the trailing edge 31B in the vicinity of the taps <NUM>. For some flight conditions, the inlet flow entering the air passage <NUM> at a high incidence with respect to the strut <NUM> may modify the expected location of the flow separation at the trailing edge 31B. Thus, a significant flow angle of attack with respect to the strut <NUM> may result in a separation delay at the trailing edge 31B. In this situation, the static pressure measurement may be skewed by a dynamic pressure contribution. The edge contour <NUM> may counter this effect by forcing the flow to separate at the trailing edge 31B in the vicinity of the taps <NUM>. 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 <NUM> may help prevent such uncertainty by forcing the flow detachment at the trailing edge 31B.

As described above, the strut passage <NUM> is in fluid communication with the taps <NUM> on the trailing edge 31B. This fluid communication arrangement may take different forms. For example, and referring to <FIG>, the strut <NUM> includes multiple measurement tap passages <NUM> that extend within the strut <NUM> and which are not exposed to the flow of air over the external surfaces of the strut <NUM>. The measurement tap passages <NUM> may be formed by any suitable technique that removes material from within the strut <NUM>, such as drilling, milling or etching. Each measurement tap passage <NUM> extends within the strut <NUM> in a direction toward the trailing edge 31B and one or more of the taps <NUM>, thereby fluid linking those taps <NUM> to the strut passage <NUM> and to the pressure sensor <NUM>. Referring to <FIG>, the diameter of the measurement tap passages <NUM> is less than the diameter of the strut passage <NUM>. The measurement tap passages <NUM> thus establish fluid communication between the taps <NUM> on the trailing edge 31B and the pressure sensor <NUM>, via the strut passage <NUM>.

The fluid communication between the strut passage <NUM> and the taps <NUM> takes another form in an alternate embodiment. In this alternate embodiment, the strut passage <NUM> is composed of two or more strut passages <NUM>. Each strut passage <NUM> is in fluid communication with one of the taps <NUM> via a dedicated measurement tap passage <NUM>. Each strut passage <NUM> is fluidly connected to its own pressure sensor <NUM>. Thus, in this alternate embodiment, the taps <NUM> are independent of one another, and each of the taps <NUM> has their own pressure sensor <NUM>.

The internal passages <NUM>,<NUM> of the strut <NUM> may be shaped or oriented to facilitate fluid drainage within the strut <NUM>, or to reduce or prevent fluid accumulation within the strut <NUM>. For example, and referring to <FIG>, the internal strut passage <NUM> and the measurement tap passages <NUM> slope toward the center axis <NUM> of the aircraft engine <NUM>. In the configuration of the strut <NUM> shown in <FIG>, each of the measurement tap passages <NUM> has an orientation defined by a predominant directional vector that is radial relative to the center axis <NUM>, such that any water which may be present within the measurement tap passages <NUM> would flow due to gravity toward the trailing edge 31B and out of the tap <NUM>. In the configuration of the strut <NUM> shown in <FIG>, drainage of the strut passage <NUM> is achieved by positioning the strut passage <NUM> so that it is radially outwardly of the location where it intersects each of the measurement tap passages <NUM>. In the configuration of the strut <NUM> shown in <FIG>, the strut passage <NUM> has a lowermost point 33LP that is the vertically lowest portion of the strut passage <NUM> when the engine <NUM> is mounted on the aircraft. The lowermost point 33LP is the portion of strut passage <NUM> that is radially closest to the center axis <NUM>. The lowermost point 33LP is positioned at the highest point (i.e. radially-outermost) of the lowermost measurement tap passage <NUM>. Referring to <FIG>, the end of the strut passage <NUM> opposite to the end engaged with the boss <NUM> is plugged or sealed.

In the configuration of the strut <NUM> shown in <FIG>, each of the measurement tap passages <NUM> has an orientation defined by a predominant directional vector that is radial relative to the center axis <NUM>, such that any water which may be present within the measurement tap passages <NUM> would flow due to gravity toward the trailing edge 31B and out of the tap <NUM>. In the configuration of the strut <NUM> shown in <FIG>, drainage of the strut passage <NUM> is achieved by positioning the strut passage <NUM> so that it is radially outwardly of the location where it intersects each of the measurement tap passage <NUM>. In the configuration of the strut <NUM> shown in <FIG>, the strut passage <NUM> has a lowermost point 33LP that is the vertically lowest portion of the strut passage <NUM> when the engine <NUM> is mounted on the aircraft. The lowermost point 33LP is the portion of strut passage <NUM> that is radially closest to the center axis <NUM>. The lowermost point 33LP allows any moisture within the strut passage <NUM> to accumulate at the lowermost point 33LP, and is positioned adjacent to the highest point (i.e. radially-outermost) of the lowermost measurement tap passage <NUM> such that water can flow from the lowermost point 33LP into, and then out of, the lowermost measurement tap passage <NUM>.

These configurations of the strut passage <NUM> and of the measurement tap passages <NUM> facilitate gravity drainage of moisture that may accumulate within the strut passage <NUM>, which may prevent accurate static pressure measurements at the taps <NUM> if the moisture is left undrained, or if the moisture freezes within the internal passages <NUM>,<NUM>. The internal passages <NUM>,<NUM> of the strut <NUM> may thus have any suitable angle with respect to the attitude of the aircraft on the ground to facilitate drainage.

Referring to <FIG>, the strut <NUM> has a first static pressure measurement tap 38A in fluid communication with the lowermost measurement tap passage <NUM>, and a second static pressure measurement tap 38B in fluid communication with the other measurement tap passage <NUM>. The first tap 38A is positioned adjacent to one of a root and a tip of the strut <NUM> and trailing edge 31B, such as adjacent to the root in <FIG>. The second tap 38B is positioned adjacent to the other of the root and the tip, such as adjacent to the tip of the trailing edge 31B in <FIG>. The trailing edge 31B has a span S measured between the root and the tip. The distance separating the first and second taps 38A,38B along the span S of the trailing edge 31B is greater than zero. The distance separating the first and second taps 38A,38B along the span S of the trailing edge 31B is not less than <NUM>% of the span S. Keeping such a minimum spanwise distance between the taps 38A,38B may minimize their sensitivity and improve the accuracy of averaging the measurement of static pressure at their locations. Alternatively, the distance separating the taps 38A,38B may be less than <NUM>% of the span S depending on the aerodynamics at the trailing edge 31B 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>, a length L of each measurement tap passage <NUM> is defined between a first end 37A at the strut passage <NUM> and a second end 37B at the trailing edge 31B of the strut <NUM>. A ratio of the length L over a diameter of each of the openings which form the taps <NUM> 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 <NUM>. The radius of each tap <NUM> may be sized with respect to the radius of the trailing edge 31B. For example, radius of each tap <NUM> may be between <NUM>/<NUM> to <NUM>/<NUM> of the radius of the trailing edge 31B. Having a bigger hole diameter for the tap <NUM> may make the tap <NUM> 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 <NUM> and potentially freeze. Selecting the diameter of the taps <NUM> 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 <NUM>, the strut passage <NUM>, the edge contour <NUM>, the measurement tap passages <NUM>, and the static pressure measurement taps <NUM> are described in relation to an inlet strut <NUM> in a radial air inlet <NUM>, the description and associated advantages of these features shown in <FIG> applies mutatis mutandis to a strut <NUM> present in an axial air inlet <NUM>, such as the one shown in <FIG>. Similarly, the description and associated advantages of the strut <NUM>, the strut passage <NUM>, the edge contour <NUM>, the measurement tap passages <NUM>, and the static pressure measurement taps <NUM> shown in <FIG> applies mutatis mutandis to the strut <NUM> present in the radial air inlet <NUM> of <FIG>, <FIG>.

Referring to <FIG>, 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.

Claim 1:
An aircraft engine (<NUM>), comprising:
an air inlet duct (<NUM>) extending from an inlet (<NUM>) through which air is configured to enter the aircraft engine to an outlet (24A) within the aircraft engine; and
at least one strut (<NUM>) having a leading edge (31A) and a trailing edge (31B) and extending across at least part of the air inlet duct (<NUM>);
characterised in that:
the at least one strut (<NUM>) has a strut passage (<NUM>) and a plurality of static pressure measurement taps (<NUM>) spaced apart on the trailing edge (31B) and in fluid communication with the strut passage (<NUM>); and
the trailing edge (31B) of the at least one strut (<NUM>) has a plurality of edge contours (<NUM>), each edge contour (<NUM>) of the plurality of edge contours (<NUM>) positioned at one static pressure measurement tap (38A, 38B) of the plurality static pressure measurement taps (<NUM>), each edge contour (<NUM>) of the plurality of edge contours (<NUM>) defining a contour edge wall (39W) recessed from a remainder of the trailing edge (31B).