Patent Description:
Exhaust gas temperature (EGT) probes are used to measure the temperature of turbine exhaust gases. As these probes are used in a harsh environment, they need to be properly protected and periodically inspected/replaced to ensure accurate EGT readings. However, with some engine architectures, access to the probes can be challenging and/or the probes may be exposed to adverse elements, which may negatively affect their durability.

<CIT> discloses a prior art gas turbine engine as set forth in the preamble of claim <NUM>.

<CIT> discloses a prior art exhaust duct for a gas turbine engine.

<CIT> discloses a prior art engine exhaust gas test harness.

From a first aspect, there is provided a gas turbine engine as recited in claim <NUM>.

There is also provide an aircraft power plant as recited in claim <NUM>.

Any of the above features may be combined in any combination, unless indicated otherwise in the description below.

<FIG> illustrates an aircraft power plant comprising a nacelle N housing a gas turbine engine <NUM> of a type preferably provided for use in subsonic flight, generally comprising 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, an exhaust system <NUM> through which the combustion gases exit the engine <NUM>. The engine <NUM> has a longitudinal center axis <NUM>. The engine <NUM> in <FIG> is a turboprop engine <NUM> and includes a propeller <NUM>, which provides thrust for flight and taxiing. The propeller <NUM> includes a nose cone 16A and propeller blades 16B, which rotate about the center axis <NUM> to provide thrust. It is understood that the engine <NUM> can adopt various other configurations. For instance, the engine could be configured as a turboshaft engine having an output shaft connectable to a rotatable load, such as a helicopter rotor or the like.

The engine <NUM> has an outer case <NUM> housing a central core through which gases flow and which includes most of the turbomachinery of the engine <NUM>. The illustrated engine <NUM> is a "reverse-flow" engine <NUM> because gases flow through the core from the air inlet <NUM> at a rear or aft portion of the engine <NUM>, to the exhaust system <NUM> at a front portion of the engine <NUM>. This is in contrast to "through-flow" gas turbine engines in which gases flow through the core of the engine from a front portion to a rear portion. The direction of the flow of gases through the engine <NUM> is shown in <FIG> with arrows F.

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

Still referring to <FIG>, the core of the engine <NUM> may include one or more spools. The illustrated embodiment is a two-spool engine including a low pressure (LP) spool and a high pressure (HP) spool rotatable about the center axis <NUM> to perform compression to pressurize the air received through the air inlet <NUM>, and to extract energy from the combustion gases before they exit the core via the exhaust system <NUM> at a forward end of the core. The core may include other components as well, including, but not limited to internal combustion engines (e.g. rotary engines such as Wankel engines for compounding power with a turbine of the turbine section), gearboxes, tower shafts, and bleed air outlets.

Each spool generally includes at least one component to compress the air that is part of the compressor section <NUM>, and at least one component to extract energy from the combustion gases that is part of the turbine section <NUM>. More particularly, according to the illustrated embodiment, the LP spool has an LP turbine 14a which extracts energy from the combustion gases, and an LP compressor 12a for pressurizing the air. The LP turbine 14a and the LP compressor 12a can each include one or more stages of rotors and stators, depending upon the desired engine thermodynamic cycle, for example. The LP spool further comprises an LP shaft <NUM> drivingly connecting the LP turbine 12a to the LP compressor 14a. Gears (not shown) can be provided to allow the LP compressor 14a to rotate at a different speed than the LP turbine 12a. The LP turbine 12a is also drivingly connected to the propeller <NUM> via a reduction gear box (RGB) <NUM>. The RGB <NUM> allows for the propeller <NUM> to be driven at its optimal rotational speed, which is different from the rotational speed of the LP turbine 14a.

Still referring to <FIG>, the HP spool comprises an HP turbine 14b drivingly engaged (e.g. directly connected) to a HP compressor 12b by a high pressure shaft <NUM>. Similarly to the LP turbine 14a and the LP compressor 12a, the HP turbine 14b and the HP compressor 12b can each include one or more stages of rotors and stators.

The LP compressor 12a, the HP compressor 12b, the combustor <NUM>, the HP turbine 14b and the LP turbine 14a are in serial flow communication via an annular gas path <NUM> extending through the core about the center axis <NUM>. The gas path <NUM> leads to the engine exhaust system <NUM> downstream of the turbine section <NUM>.

The exhaust system <NUM> of the engine <NUM> comprises a turbine exhaust duct <NUM> for exhausting combustion gases received from the last stage of the LP turbine 14a. According to the illustrated embodiment, the exhaust duct <NUM> is a non-axisymmetric dual port exhaust duct configured for directing combustion gases laterally on opposed sides of the outer case <NUM> of the engine <NUM> of the aircraft power plant. The dual port exhaust duct <NUM> is qualified as "non-axisymmetric" because the two exhaust ports thereof are not coaxial to the center axis <NUM> of the engine (i.e. the exhaust flow discharged from the exhaust duct is not axial, it is rather directed in a direction that diverges from the center axis <NUM>). According to at least some embodiments, the dual port exhaust duct <NUM> has a generally "Y-shaped" body including an annular central inlet conduit portion 30a extending axially around the center axis <NUM> for receiving the annular flow of combustions gases discharged from the last stage of LP turbine 14a, and first and second diverging outlet conduit portions 30b, 30c branching off laterally from the central inlet conduit portion 30a. According to some embodiments, the first and second outlet conduit portions 30b, 30c are identical.

As can be appreciated from <FIG>, the downstream end of each outlet conduit portion 30b, 30c projects outwardly of the engine outer case <NUM> into an air space or gap G between the outer case <NUM> and the nacelle N. As best shown in <FIG>, with reference to the first outlet conduit portion 30b, each outlet conduit portion 30b, 30c terminates into an exhaust port including an annular outer flange <NUM> surrounding a central exhaust port opening <NUM>. The central exhaust port has a central axis A which is oriented to intersect the center axis <NUM>. According to the illustrated embodiment, the axis A has a main radial component and a secondary (i.e. smaller) axial component relative to the center axis <NUM>. Stated differently, the exhaust port is oriented to direct the combustion gases mainly in a radially outward direction. According to some embodiments, the exhaust port opening <NUM> and surrounding annular flange <NUM> are circular. However, it is understood that other geometries are contemplated as well (e.g. oval).

Referring back to <FIG>, it can be seen that first and second tailpipes TP are respectively detachably mounted to the annular flange <NUM> of the first and second outlet conduit portions 30b, 30c. The tailpipes TP can be provided with an annular flange complementary to the annular flange <NUM> of the exhaust duct <NUM>. The tailpipes TP extend outwardly from the nacelle N and curve towards a rearward direction to discharge the combustion gases received from the exhaust duct <NUM> into the surrounding environment on opposed sides of the nacelle N and with a generally rearward component.

The temperature of the exhaust gases may be used to measure the performance of the engine <NUM> and to provide an indication of the rate of deterioration of gas turbine engine components. Indeed, the exhaust gas temperature (EGT) is an indicator of engine status, which may be used to measure and control operational and functional characteristics of the engine <NUM>.

Accurate measurement of the EGT level is thus important. To accurately measure exhaust gas temperatures, it is necessary to minimize degradation of the EGT measurement system over time. Therefore, it is desirable to periodically inspect the EGT measurement system and to properly shield the same from the harsh environment in which it is used.

According to the illustrated embodiment, the EGT measurement system comprises first and second sets of operatively interconnected EGT probes <NUM>, <NUM>. The EGT probes <NUM>, <NUM> are positioned outside the outer case <NUM> of the engine <NUM> in the air space/gap G between the nacelle N and the outer case <NUM>. More particularly, the first set of EGT probes <NUM> is provided at the exhaust port of the first outlet conduit portion 30b, and the second set of EGT probes <NUM> is provided at the exhaust port of the second outlet conduit portion 30c. By so instrumenting the two exhaust ports of the exhaust duct <NUM> and positioning the EGT probes <NUM>, <NUM> outside the engine outer case <NUM>, access to the probes <NUM>, <NUM> can be facilitated for maintenance and inspection purposes while still providing for accurate monitoring of the engine exhaust gas temperature. Indeed, as opposed to EGT probes positioned in the turbine section of the engine, it is not necessary to split open the gas generator/hot case section of the engine to access the probes.

As best shown in <FIG>, the EGT probes <NUM>, <NUM> may be distributed in predetermined positions around the exhaust port of each outlet conduit portion 30b, 30c of the exhaust duct <NUM> and inserted to a predetermined depth into the duct to establish the temperature of the combustion gases as they exit the exhaust duct <NUM>. According to an embodiment, six probes are uniformly circumferentially distributed around the exhaust port of each outlet conduit portion 30b, 30c. The exhaust gas temperature profile for a particular engine may be determined by using a plurality of probes or thermocouple elements so arranged around each exhaust port at various penetration depths.

The EGT probes <NUM>, <NUM> of a same set are electrically interconnected by a wiring harness (not shown) outside of the exhaust duct <NUM>. The probes <NUM>, <NUM> and the associated harness at the end of each outlet conduit portions 30b, 30c are protected by a shield <NUM>. As best shown in <FIG>, each shield <NUM> has a cap-shaped body including a cylindrical skirt 60a extending from an outer periphery of end wall forming an annular inward flange 60b around a central opening 60c. The flange 60b is complementary to the annular flange <NUM> at the end of each outlet conduit portion 30b, 30c of the exhaust duct <NUM> and is adapted to be detachably connected thereto such as by bolting or the like. According to the embodiment shown in <FIG>, the flange 60b of the shields <NUM> are sandwiched between the flanges of the exhaust duct <NUM> and the tailpipes TP. That is the shields <NUM> are mounted at the interface between the exhaust duct <NUM> and the tailpipes TP. In this way, access to the probes <NUM>, <NUM> can be readily provided by opening the nacelle N, detaching the tailpipes TP from the dual port exhaust duct <NUM> and then removing the shields <NUM> from the exhaust ports of the exhaust duct <NUM>. Once the probes <NUM>, <NUM> have been inspected, replace or reinstalled, the shields <NUM> can be bolted back to the exhaust ends of the exhaust duct <NUM> in order to protect the probes <NUM>, <NUM>. As best shown in <FIG>, once installed in position, the skirt 60a of each shield <NUM> surrounds the associated probes <NUM>, <NUM> all around each exhaust end of the exhaust duct <NUM>. The central opening 60c of the shield <NUM> may be sized to be slightly larger than the exhaust port opening <NUM> so as to not interfere with the flow of exhaust gases. The central opening 60c of the shield is centered relative to axis A once the shield <NUM> has been properly secured to the flange <NUM> at each exhaust end of the exhaust duct <NUM>. The skirt 60a is configured to form an annulus around each exhaust end of the exhaust duct <NUM>. The head of the probes <NUM>, <NUM> and the wiring harness are accommodated in this annulus. The shields <NUM> thus form a physical barrier to protect the probes <NUM>, <NUM> and associated wiring in the air space/gap G from tool, parts or fluid surrounding the engine outer case <NUM>. In addition to providing a physical protection, the shields <NUM> provide thermal shielding to the probes <NUM>, <NUM>.

According to some embodiments, the shields <NUM> contribute to improve the EGT probes <NUM>, <NUM> functionality and durability by decreasing the risk of probes or wire damages.

According to some embodiments, each shield <NUM> has a unitary body made from Inconel <NUM> or other suitable sheet metal material pressed or stamped into an inverted cup-shaped body.

Claim 1:
A gas turbine engine (<NUM>) comprising:
an outer case (<NUM>);
a core inside the outer case (<NUM>), the core including:
a combustor (<NUM>),
at least one spool mounted for rotation about a central axis (<NUM>), the at least one spool including a compressor (<NUM>) and a turbine (<NUM>),
an annular gas path (<NUM>) extending between the compressor (<NUM>) and the turbine (<NUM>) about the central axis (<NUM>); and
a turbine exhaust duct (<NUM>) extending from the annular gas path (<NUM>) downstream of the turbine (<NUM>) in a direction away from the central axis (<NUM>), the turbine exhaust duct (<NUM>) having a downstream end portion (30b; 30c) projecting outwardly from the outer case (<NUM>), the downstream end portion (30b; 30c) instrumented with exhaust gas temperature (EGT) probes (<NUM>; <NUM>),
characterised in that:
the EGT probes (<NUM>; <NUM>) are electrically interconnected by a wiring harness outside the turbine exhaust duct (<NUM>) and are distributed around the downstream end portion (30b; 30c) of the turbine exhaust duct (<NUM>) outside the outer case (<NUM>) of the gas turbine engine (<NUM>); and
the downstream end portion (30b; 30c) of the turbine exhaust duct (<NUM>) defines an exhaust port opening (<NUM>) surrounded by an annular flange (<NUM>), and wherein a shield (<NUM>) is detachably mounted to the annular flange (<NUM>), the shield (<NUM>) having an annular skirt (60a) surrounding the EGT probes (<NUM>; <NUM>) around the downstream end portion (30b; 30c) of the turbine exhaust duct (<NUM>), the annular skirt (60a) forming an annulus around the downstream end portion (30b; 30c) of the turbine exhaust duct (<NUM>), the wiring harness accommodated in the annulus.