TURBINE ENGINE WITH INDUCED PRE-SWIRL AT THE COMPRESSOR INLET

Abstract: A jet flap construction for the compressor of an aircraft gas turbine engine provided with a bleed means for bleeding air interstage of the compressor and communicating with the hollowed-out struts in the airflow path to the compressor. The hollow struts are each provided with one or more rows of discrete openings extending from the shroud to the hub on the strut to form an aligned series of jets for discharging the pressurized bleed air and forming a jet flap, thereby providing pre-swirl at the compressor inlet.

Description: The present invention relates to a gas turbine engine, particulary to a jet flap arrangement for producing a pre-swirl at the compressor inlet. Attempts have been made in the past to overcome the problem of engine surge at off design modes in gas turbine engines while producing pre-swirl in the compressor inlet by bleeding air interstage of the compressor and advantageously using this bleed air to form a jet flap on hollow struts in the compressor inlet such as described in U.S. Patent 4, 222, 703, Schaum et al, 1980. The Schaum et al patent describes hollow struts in the compressor inlet preceding the first compressor stage communicating with a passage and a bleed valve for bleeding air from interstage of the compressor. Each hollow strut has a slot near the trailing edge of the strut for producing the jet flap. The strut extends between the shroud and the hub in that portion of the engine thereof. It has been found that the inertia of the bleed air moving in the strut from the shroud to the hub will cause the air distribution at the slot forming the jet flap to be unequal and for the highest pressure to be nearest the hub, thereby affecting the pre-swirl at the compressor inlet. It is an aim of the present invention to provide a gas turbine engine with an improved jet flap arrangement in order to create a more efficient compressor pre-swirl to enhance engine stability and performance at low speed off-design conditions while avoiding the draw backs due to the inertia of the bleed air which causes unequal air distribution affecting the pre-swirl at the compressor inlet. It is another aim of the invention to provide a method for providing a predetermined pre-swirl in different radial sectors of the inlet of the compressor of such a gas turbine engine. The gas turbine engine as claimed, having an axial compressor with induced pre-swirl at the inlet of the compressor, includes a shroud wall and a hub wall defining an airflow path to the compressor inlet having a radial portion and an axial portion, and a plurality of hollow strut members extending across the airflow path between the shroud wall and the hub wall of the engine substantially within said radial portion upstream of the compressor. The hollow strut members communicate with an air inlet in the shroud wall for communicating with a source of pressurized gas flow. Each hollow strut member has a leading edge and a trailing edge relative to the airflow path and a passage provided near the trailing edge such that the gas flowing into each of said hollow strut members will be emitted by said passage at an angle to the airflow, thereby creating a pre-swirl in the airflow at the inlet of the compressor. According to the invention, said passage comprises at least one row of a plurality of discrete openings lying on a axis extending from said shroud wall to said hub wall and each defining an individual nozzle to form a jet, each said opening having a selected location, size and angle adapted to provide a predetermined pre-swirl angle in each specific radial sector of the airflow path at the inlet of the compressor. The selective location, size and angle of the discrete openings will determine the distribution of pre-swirl in the airflow path in front of the compressor. In other words, although the provision of a plurality of openings in one or more rows in the hollow strut across the airflow path may have the effect of equalizing the air distribution across the airflow path to form a more effective jet flap than that described in the Schaum et al patent, the actual determination of the size of the different discrete openings will affect the actual jet flap angle and thus the pre-swirl of the airflow at different radial distances across the airflow path in front of the compressor. For instance, the angle of pre-swirl required near the tip of the first stage compressor rotor may be different from the angle of pre-swirl of the airflow near the hub of the rotor, as a result of the twist of the rotor blades on the rotor, keeping in mind that it is a purpose of pre-swirl to reduce the angle of incidence of the airflow to the rotor blades and thereby improve performance and compressor stability. Thus, in any predetermined engine design, a bleed air jet flap can be designed by providing on the hollow struts within the compressor inlet a pattern of discrete holes which have been determined by the angle of pre-swirl required at the face of the compressor in different sectors across the airflow path determined by the compressor inlet. In other words, given an engine compressor section having a circumferential inlet determined by an airflow path leading to an axial compressor having a first stage rotor wherein the airflow path is defined by a hub side wall and a shroud side wall and wherein radially extending hollow struts are spaced apart in the inlet between the hub and shroud walls, the gradient of the pre-swirl angle relative to the rotor can be varied radially from the hub of the rotor to the tips of the rotor. It is necessary, therefore, to determine the required angle of pre-swirl in the airflow relative to the rotor at different radial sectors of the airflow path in front of the compressor and to translate these different angular profiles to the pre-swirl ultimately to the pattern of discrete openings provided in the hollow struts which form the jet flap. In other words, sectors across the height of the strut corresponding to the radial sectors at the face of the compressor can be provided with discrete openings which can be designed in terms of their size, location, and angle to provide the required pre-swirl angle in the different radial sectors at the face of the compressor within certain parameters. More specifically, one or more rows of openings may be provided in the wall of the hollow strut extending across the airstream path from the shroud to the hub. The openings may be larger near the shroud and be of gradually smaller size towards the hub. The advantages of such a construction, that is, of having a row or plurality of rows of openings across the airflow path from the shroud to the hub, is to provide a better distribution of air forming the jet flap, thereby improving the pre-swirl distribution at the entry of the compressor. In another embodiment, the closely spaced-apart openings may be in the form of an elongated tapered slot with the larger width of the slot near the shroud extending towards the hub and a series of partitions extending across the slot defining the discrete openings. The partitions could be in the form of an insert having a common base member with spaced-apart partition walls extending at 90° from the base member, the insert being adapted to be fitted into the slot. Having thus generally described the nature of the invention, reference will now be made to the accompanying drawings, showing by way of illustration, a preferred embodiment thereof, and in which : Figure 1 is a fragmentary axial cross-section of a detail of a typical turbine engine compressor; Figure 2 is an enlarged fragmentary longitudinal cross-section of a detail shown in Figure 1 but modified to show a different embodiment thereof; Figure 3 is a vertical cross-section taken along line 3-3 of figure 2; Figure 4 is an enlarged fragmentary longitudinal cross-section similar to Figure 2 but showing yet another embodiment thereof; Figure 5 is a fragmentary enlarged view of a detail shown in Figure 4; Figure 6 is a graph showing the pre-swirl angle profile of the airflow at the compressor face in degrees against the radial distance from the hub of the rotor; and Figure 7 is a graph showing the ratio of the pre-swirl angle of the airflow at the compressor face over the angle of inclination of the jet nozzle against the ratio nozzle jet momentum to the airflow momentum ratio at the strut. Referring now to the drawings and particulary to Figure 1, there is shown a fragment of a typical turbine engine 10 illustrating the compressor section. The engine 10 includes an outer casing 12 surrounding a multi-stage, single shaft compressor 14. As part of the compressor, a centrifugal impeller 16 is mounted on the same shaft (not shown). A diffuser 18 is concentric with the outlet of the impeller 16. An annular chamber 20, which might be referred to as a plenum chamber, surrounds to compressor 14. One or more bleed valves 22 will control the flow of air from the plenum chamber 20 to the hollow struts 24. A plurality of radially and axially extending engine support struts 24 are normally provided in the inlet case 25. The struts 24 in the present invention are hollow as will be described later. The main airflow path 50 is defined between the shroud 46 and the hub 48. The compressor 14 in this particular embodiment includes a leading rotor section 26 having a plurality of blades mounted to the shaft (not shown) with subsequent rotor stages represented by rotor blades 30 and 34. Interstage of the rotor blades 26, 30 and 34 are the stator vane stages 28 and 32. Finally, between the rotor 34 and impeller 16 is a stage of stator vanes 36. A plurality of bleed ports 38 is provided in the present embodiment between the stator vane stage 36 and the impeller 16 to bleed air from the compressor into the plenum 20, particulary at low power modes. The bleed ports 38 could be provided between other stages of the compressor but preferably not at the exit of the compressor that is beyond the diffuser 18 since considerable work has been done to the air by this stage. It has been noted that the bleed port 38 can also be provided at the impeller shroud. A valve seat 40 is provided with each valve 22. The valve 22 and valve seat 40 are substantially as described in U.S. patent Nos. 3, 360, 189 and 4, 222, 703. In the former patent, the valve controls bleed air from the compressor to allow the bleed air to discharge overboard the engine. In the latter patent, the valve chamber communicates with a plurality of conduits which pass the pressurized bleed air to each individual hollowed-out support strut via a communicating volute to equalize flow distribution. Similarly, in the present embodiment as shown in Figure 1, the conduits 42 pass the pressurized bleed air from the valve chamber 44 and plenum 20 through the communicating volute 60 into the hollowed-out support struts 24. The support strut 24, as shown in Figure 3, has a cavity 56, communicating with the volute 60. In the embodiment shown in Figures 1, 2 and 3, different size bores or openings 58 are drilled in the side wall of the strut 24. In the embodiment shown in Figure 1, the openings 58 have gradually decreasing diameters from the shroud 46 to the hub 48. On the other hand, it is understood that these openings may be of constant diameter or inversely decreasing depending on the surge problems of a specific engine installed. Each of the openings 58 is aligned in a row having an axis substantially parallel to the trailing edge 54 of the strut 24. In the present embodiment, a line tangent to the bottom of the circular openings 58 is parallel to the trailing edge 54. In Figure 1, only one row of openings 58 is provided, while in Figure 2, two rows of openings 58 are illustrated with the openings staggered as shown. The openings 58 are drilled at an angle τ to the apparent median plane of the strut 24. In tests, the angles were selected at between 30° and 60° from the plane. In operation, when the engine is idling or at low speed, the valve 22 is operated so as to retract from the valve seat 40 and allow air to bleed from the ports 38 to the plenum 20 through the valve chamber 44, the conduits 42, the volute 60 and into the hollow cavities 56 of the struts 24 to then be discharged under pressure through the openings 58 forming a jet flap in order to provide pre-swirl to the air entering the engine in the inlet 25. The concept of providing pre-swirl in the airflow by means of a jet flap is well described in U.S.Patent 4, 222, 703. The present arrangement of the discrete openings 58 allows for better distribution of the pressurized jet flap across the main airflow passage. Since the momentum of the air entering the cavity 56 from the shroud side of the strut 24 is to move towards the hub, the openings 58 nearest the hub are smaller than the openings near the shroud in order to enhance the jet flap uniformity across the airflow path. EXAMPLE 1 The width of air path from the shroud to the hub at the level of the jet flap - 3 in.(7,62 cm) Number of rows - 1 Number of openings - 8 Diameter of opening nearest shroud - 0.386 in.(0,98 cm) Diameter of opening nearest hub - 0.136 in.(0.35 cm) Total area of openings - 0.46 in.²(2,97 cm²) Spacing between openings in straight line -0.12 in.(0,30 cm) The line tangent to the bottom of the openings is parallel to the strut trailing edge and is located about 0.25 in. upstream from the trailing edge of the strut. Angle of bore of openings from median plane of strut - 45° EXAMPLE 2 The width of air path from the shroud to the hub at the level of the jet flap - 1.8 in.(4,57 cm) Number of rows - 2 Number of openings - 9 Diameter of opening nearest shroud - 0.400 in.(1,02 cm) Diameter of opening nearest hub - o.160 in.(0,41 cm) Total area of openings - 0.493 in.²(3,18 cm²) Angle of bore of openings from median plane of strut - 45° In the case of two or more rows of openings, advantages were found over a single row. For instance, more pressure bleed air is used to form the jet flap and thus almost all bleed air can be used rather than dumping the bleed air overboard. bleed air already has energy, having been bled downstream of the compressor. Multiple rows with the openings staggered provide better distribution of the jet flap air. The downward angle τ of the bore of the openings is important to the pre-swirl design. This angle may be varied. So far, tests have shown that the angle can be anywhere between 30° and 60°. The embodiment shown in Figure 4 shows a strut 24b with a leading edge 52b and a trailing edge 54b. The shroud 46 and hub 48 are also represented. In this embodiment, as shown in Figures 4 and 5, a conventional jet flap slot 58b is provided in the strut 24b. The discrete openings are defined by the provision of an insert 55 having a base 57 and several upstanding partition walls 59. The partition walls 59 extend across the entire width of the slot 58b in order to form the discrete openings and thus direct the flow of pressurized gas forming the jet flap. In addition to obtaining a better distribution of bleed air forming the jet flap across the airflow path 50 as defined by the shroud 46 and the hub 48, a varied angular profile of the pre-swirl at the compressor rotor stage 26 can be provided as will now be described. As is well known, the blades of the first stage rotor 26 will have varying degrees of blade leading edged angle from the root of the blade at the hub to the tip of the blade. The angle of incidence, therefore, changes along the profile of the leading edge of the rotor radially from the hub to the tip (the tip being closest to the shroud). In order to predict and properly provide for the variable angular profile of the pre-swirl at the face of the compressor, a pattern of angles α at the compressor face in different radial sectors of the airflow path is first determined at a given rpm. Then, a model of the airflow path is made at the level of the strut 24, that is, near the trailing edge 54 thereof, in order to determine the m o (the mass airflow rate lb/s) in the different radial sectors determined by the streamlined pattern in those sectors. Likewise, V o' (velocity of the airflow ft/s) is also determined. M o V o is the airflow momentum in the airflow path upstream of the openings 58. Referring now to Figure 7, it can be seen that, given the angle α required in a particular radial sector and knowing m o v o in that sector, m J (mass flow rate of the jet from the opening 58) and V J' (velocity of the jet at the nozzle 58) can be determined, keeping in mind the angle τ which can vary from 30 to 60°. It has been found that the angle τ is preferably at 45°. Thus having decided the jet momentum m j v j in each sector, it is a simple matter to determine the size or area of the opening and the distribution of the opening in a particular sector, that is, whether, given structural aerodynamic factors, several rows of openings are required in the strut or whether larger or smaller openings are required. Figure 6 shows a particular selection of an angular profile of the pre-swirl at the compressor face or at the first stage rotor between the hub and the shroud as a result of providing two rows of holes and having designed the size of the holes to meet the angular profile required at the compressor face following the above methods. The ratio m J v j /m o v o' , as shown in Figure 7 is, in fact, the momentum ratio of the jet over the air stream. As can be seen from Figure 7, it is this ratio which will affect the angular profile of the pre-swirl at the compressor face.

Claims: Gas turbine engine (10) having an axial compressor (14) with induced pre-swirl at the inlet of the compressor , the engine including a shroud wall (46) and a hub wall (48) defining an airflow path to the compressor inlet having a radial portion and an axial portion , a plurality of hollow strut members (24) extending across the airflow path between the shroud wall (46) and the hub wall (48) of the engine substantially within said radial portion upstream of the compressor, the hollow strut members (24) communicating with an air inlet (38) in the shroud wall (46) for communicating with a source of pressurized gas flow, each hollow strut member (24) having a leading edge (52) and a trailing edge (54) relative to the airflow path and a passage (58) provided near the trailing edge such that the gas flowing into each of said hollow strut members (24) will be emitted by said passage at an angle to the airflow , thereby creating a pre-swirl in the airflow at the inlet of the compressor , characterized in that said passage comprises at least one row of a plurality of discrete openings (58) lying on an axis extending from said shroud wall (46) to said hub wall (48) and each defining an individual nozzle to form a jet, each said opening having a selected location, size and angle adapted to provide a predetermined pre-swirl angle (α) in each specific radial sector of the airflow path at the inlet of the compressor. Gas turbine engine as defined in claim 1, wherein the combined area of the discrete openings (58) decreases gradually from the shroud wall (46) to the hub wall (48). Gas turbine engine as defined in claim 1 or 2, wherein the discrete openings (58) are in the form of bores in the wall of the hollow strut member (24a), extending at an acute angle (τ) from the median plane of the hollow member (24) in a direction downstream of the hollow strut member (24). Gas turbine engine as defined in claim 3, wherein the acute angle (τ) of the bores (58) is between 30 o and 60 o to the median plane of the hollow strut member (24). Gas turbine engine as defined in claim 4, wherein the angle of the bores (58) is 45 o to the median plane of the hollow strut member (24). Gas turbine engine as defined in claim 2, wherein the openings (58) are circular bores, and each bore has a diameter gradually decreasing from the bore (58) closest to the shroud wall (46) to the bore closest to the hub wall (48). Gas turbine engine as defined in anyone of the preceding claims, wherein there are at least two rows of discrete openings (58) in the wall of the hollow strut member (24), the openings (58) in one row being staggered with the openings (58) in the other row so as to provide a uniform distribution of pressurized gas to form a jet flap. Gas turbine engine as defined in anyone of the preceding claims, wherein the discrete openings (58) are in the form of an elongated slot (58b) defined in the wall of the hollow strut member (24b) with partitions (59) extending across the width of the slot (58b) to define separate and distinct discrete openings (58) successively along the slot (58b). A method of providing a predetermined pre-swirl in different radial sectors of the inlet of the compressor of a gas turbine engine as defined in anyone of the preceding claims, characterized in that it comprises the steps of determining the required angle (α) of pre-swirl in a predetermined radial sector of the inlet of the compressor, determining the mass airflow rate (m o ) and the airflow velocity (V o ) upstream of the hollow strut member in a corresponding radial sector and calculating the corresponding airflow momentum (m o V o ), determining the jet flap momentum (m j V j ) required in said sector, where m j is the mass flow rate of the jet flap and V j is the velocity of the jet flap in said sector, and determining the angle of inclination (τ), the size and the location of the discrete openings required in the hollow strut member to form in said sector individual jets providing said jet flap momentum (m j V j ). The method as defined in claim 9, wherein the angle of inclination (τ) of the discrete openings is between 30° and 60° to the median plane of the hollow strut member.

Inventors: HUBINETTE, SVEIN, LACHANCE, ROGER, LETOURNEAU, JEAN, OZARAPOGLU, VASIL
Applicants: PRATT & WHITNEY CANADA INC.
Application year: 1986
Main classification: F02C7
IPC codes: F02C9, F04D29, F04D27