RU2397351C2 - Turbine engine (versions) - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: proposed engine comprises axial-flow compressor, combustion chamber, turbine and nozzle. One, several or all compressor stages comprise vanes that feature, in plan, triangular shape and/or height equal to boundary layer thickness. Via hollow vanes of stator or compressor or blade rim, pressure is equalised over compressor stage section by bypassing air from peripheral zone of increased pressure into circum-rotor zone with lower pressure. Hollow vane peripheral parts have holes or slots arranged on front edge or vane profile concave side, while circum-rotor part has holes or slots arranged on rear edge of vane profile convex side. Compressor and/or turbine vanes feature smoothly decreasing pitch on periphery and/or circum-rotor parts on the length that corresponds to boundary layer thickness. Turbine engine has additional rim or additional stage with vanes with height equal to boundary layer thickness.

EFFECT: higher anti-surge properties, engine response and reliability.

8 cl, 3 dwg

Description

The invention relates to turbojet engines (turbojet engines) and gas turbine engines (GTE), as well as gas axial compressors and steam turbines.

Known turbojet engines, see, for example, US Pat. Russia 2067683 or 1763695, consisting of an axial compressor, combustion chamber, turbine, afterburner and nozzle. The engine compressor is prone to surge during transient conditions. Presumably, one of the reasons accelerating the development of surging is the pressure difference at the periphery of the stage, where the peripheral speed of the blades is greater, and near the rotor, where this speed is lower, that is, they operate more efficiently on the periphery and the pressure created by them is greater, and much more. Perhaps, as a result of this, a radial flow of air into the zone with lower pressure arises, passing into the oncoming flow in the near-rotor zone, which accelerates the stall flow around the blades.

The second possible reason for accelerating surge is the presence of a boundary layer at the surface of the rotor and near the surface of the outer shell of the engine, where the flow rate is lower.

All the following technical solutions are aimed at improving the anti-surge properties, improving the throttle response, increasing reliability, and invention 4, in addition, also obtaining a new quality of the engine - its ability to work in the boundary layer at the rear end of the fuselage.

All turbo engine options include a compressor, a combustion chamber, a turbine, a jet nozzle, and possibly a second circuit.

Invention 1. Since we cannot increase the efficiency of the blades in the near-rotor (basal) zone, to reduce the mentioned pressure difference, it is necessary to reduce the efficiency of the blades on the periphery. To equalize the pressure over the cross section of the compressor stage, one, several or all stages of the compressor have vanes that taper to the periphery, in particular to zero and / or have a smaller pitch on the periphery, and / or have a lower height, counting from the rotor, i.e. a smaller diameter crown than the diameter of the stator, but having a height not less than the thickness of the boundary layer (hereinafter “step” is the distance by which the blade is screwed in the longitudinal direction for one revolution, and the height of the blade is the difference between the radius of the crown and the radius of the rotor) .

For example, the engine may have an additional lightweight stage with blades having a plan in the expanded form of a triangle, that is, tapering to the periphery almost to zero. The optimal place for such a stage is closer to the compressor inlet, that is, the second or third.

Or, the engine can have all stages with blades having a trapezoidal plan in expanded form. In this case, the engine will also have one or two more compressor stages than usual. But since the blades will be lighter, the total weight of the engine will not change or even decrease.

And the reliability, throttle response and anti-surge properties of this engine will increase significantly.

Or the engine can have one or two stages with blades having a height equal to 60, 50, 40% of the maximum possible in a given diametrical section.

Invention 2. It is possible to equalize the pressure over the cross section of the compressor stage by passing air from the peripheral zone of increased pressure to the near-rotor zone with reduced pressure along the hollow blades of the stationary guide vane or along the hollow blades of the compressor rim. To do this, the hollow vanes of the guide vanes or compressor crowns, through which the pressure is equalized over the section of the compressor stage, passing air from the peripheral pressure zone to the low-pressure rotor zone, have holes or slots on the peripheral part located on the leading edge or on the concave side of the profile the blades, and on the near-rotor part there are holes or slots located on the trailing edge or on the convex side of the blade profile.

In this case, when considering the feasibility of such a bypass through the hollow blades of the crown, one should take into account the centrifugal force acting on the air inside the rotating blade. It may turn out to be greater than the pressure difference at the periphery of the crown and near the rotor.

The invention 3. To prevent the initiation of flow stall in the zone of the boundary layer moving at a lower speed, the compressor blades and / or turbines have a smoothly decreasing step on the periphery and / or near-rotor parts at a length corresponding to the thickness of the boundary layer so as not to provoke supercritical flow around the profile .

Invention 4. For the same purpose, the engine has an additional crown or additional step with blades, the height of which is equal to or greater than the thickness of the boundary layer. That is, the near-rotor boundary layer is blown away, and the pressure in the near-rotor zone rises. The best place for such a crown is in front of the first stage of the compressor. If the compressor does not have an input guide vane, then the direction of rotation of the additional crown should be opposite to the direction of rotation of the first stage of the compressor.

Since the power required to drive such a crown is relatively small, it is not always advisable to bring it into rotation by the main turbine. A turbo engine can consist of two engines of different power, and a lower power engine rotates an additional crown. Such an engine can be a small internal combustion engine (ICE), for example a star-shaped one, which through the gearbox and friction clutch can act as a starter for the main engine. The use of such an engine is especially appropriate when the engine is located at the rear of the fuselage. In this case, the engine with its rotor can adjoin the rear of the fuselage.

This promises benefits in the layout, aerodynamics and engine maintenance, and also protects the engine well from birds and shelling in front. The effectiveness of this additional crown or step can be increased by releasing its own boundary layer. For this, the engine has an additional crown or stage, the rotor diameter of which is less than the inner diameter of the first stage of the compressor, and between them there is an annular gap with the possibility of air exhaust. For example, into an ejector nozzle device through a hollow shaft.

Figure 1 is a simplified sectional view of a conventional single-circuit turbojet engine containing an axial compressor 1, consisting of oppositely rotating crowns 2, a combustion chamber 3, a turbine 4, an afterburner 5, and a nozzle 6. There is an incomplete stage for equalizing the pressure across the compressor cross section 7, the crowns of which have a plan view close to a triangle.

In front of the compressor there is an additional crown 8, driven by a separate engine 9 located in the fuselage 10. Moreover, the fuselage is adjacent to the multi-shaft rotor 11. Between the additional crown 8 and the first crown 2 of the compressor there is an annular gap 12 into which the boundary layer of the additional crown is removed and ejected in the ejector 13.

An oxidizing agent and an evaporating liquid may be supplied to the combustion chamber, and an oxidizing agent (shown by arrows) may be supplied to the afterburner.

Figure 2 shows a fragment of an engine compressor consisting of a multi-shaft rotor 11, a stator 1a, full-profile vanes 2, hollow vanes of the guide apparatus 14 and an auxiliary shell 15 fixed on them. Curved arrows show the passage of air through the cavity of the vanes 14 into the near-rotor zone to blow off the boundary layer.

Figure 3 shows an exemplary view of the aircraft with the proposed engine. The aircraft consists of a fuselage 10 with a wing 16 and two longitudinal pylons 17, which are the outer side members. At the rear, the pylons go into keels 17. On these pylons and on two wing consoles 16, an engine 19 is mounted with an adjustable input device 20. The compressor rotor, as shown in figure 1, is adjacent to the rear of the fuselage. There is a front horizontal plumage 21.

The engine works as follows: a separate engine 9 accelerates the boundary layer to the speed of the main stream or slightly more. Opposite rotating crowns 2 compress the air, and the incomplete crown 7 creates an increased pressure in the near-rotor zone, due to which the peripheral pressure drop created by the main crowns 2 is leveled.

The engine of FIG. 2 operates as follows: oppositely rotating vanes 2 compress the air. In this case, due to the greater peripheral speed of the blades at the stator 1a, a zone of increased pressure is formed. Partial-profile blades 7, having a height of 50% of the possible and having a trapezoidal profile in plan, create pressure in the near-rotor zone, leveling the specified pressure drop. To prevent the stall of the flow, air from the front near the part of the blades 14 is discharged into the rear side of the rotor part.

The plane of FIG. 3 operates as follows: air from the rear of the fuselage 10 before entering the compressor 1 of the engine 19 enters an additional crown 8 (see FIG. 1), where the speed of the boundary layer is leveled and brought to the speed of the main stream or higher. This arrangement of the engine allows air to be supplied to the engine without special air ducts and trays for removing the boundary layer, which in general will increase traction and reduce aerodynamic drag.

Claims (8)

1. A turbo engine containing an axial compressor, a combustion chamber, a turbine and a nozzle, characterized in that one, several or all stages of the compressor have blades having a triangular plan in plan view and / or having a height equal to the thickness of the boundary layer. 2. A turbo engine containing an axial compressor, a combustion chamber, a turbine and a nozzle, characterized in that the hollow vanes of the guide vanes or compressor crowns through which pressure is equalized over the compressor stage cross-section, passing air from the peripheral pressure zone to the low-pressure rotor zone, have holes or slots on the peripheral part located on the leading edge or on the concave side of the blade profile, and on the rotor part have holes or slots located on the trailing edge if on the convex side of the blade profile. 3. A turbo engine containing an axial compressor, a combustion chamber, a turbine and a nozzle, characterized in that the compressor blades and / or turbines have a smoothly decreasing step on the periphery and / or near-rotor parts at a length corresponding to the thickness of the boundary layer. 4. A turbo engine containing an axial compressor, a combustion chamber, a turbine and a nozzle, characterized in that it has an additional crown or additional stage with blades, the height of which is equal to the thickness of the boundary layer. 5. The turbo engine according to claim 4, characterized in that it consists of two engines of different power, and the engine of lower power rotates an additional crown. 6. The turbo engine according to claim 5, characterized in that the engine of lower power is a star-shaped internal combustion engine, which together with the gearbox and friction clutch is the starter of the main engine. 7. The turbo engine according to claim 4, characterized in that it has an additional crown or stage, the rotor diameter of which is less than the inner diameter of the first compressor stage, and between them there is an annular gap with the possibility of air exhaust. 8. The turbo engine according to claim 7, characterized in that the air is discharged into the ejector device of the nozzle through the hollow shaft.

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