JPH0154524B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to cooling air passages for high temperature turbines.

In recent years, gas turbine engines have employed nozzles that accelerate cooling air in the direction of rotation of the turbine to facilitate entry of the cooling air into the rotating shaft. A nozzle directs cooling air into a round hole drilled in the shaft wall. Before reaching the turbine blades, the cooling air passes through these holes in the shaft and then moves radially outward to the hot turbine blades.

Adopting this improvement has a reasonable effect, but
Further improvement is expected. The analysis showed that after leaving the turbine nozzle, the cooling air has a very high flow velocity, which is faster than the rotational speed of the turbine shaft. This means that the cooling air has a velocity tangential to the shaft. When the air is decelerated and passes through the shaft hole, a large pressure loss occurs. This represents an irrecoverable energy loss. Also, when the cooling air flows to a smaller radius, such as when passing under the turbine disk, its tangential velocity increases further and may actually cause acoustic resonance. More recently, turbine engines have been provided with flat radial vanes inside the turbine shaft to eliminate residual tangential flow velocity of cooling air. Although this measure was successful in eliminating acoustic resonance, it further increased aerodynamic losses in the cooling airflow system. These potential problems can be avoided by removing excess tangential flow velocity of the cooling air before it enters the turbine shaft.

It is therefore an object of the present invention to transfer non-rotating cooling air to the rotation of a gas turbine engine without significant pressure losses, yet with lower resulting cooling air temperatures and reduced cooling air flow requirements. The purpose is to introduce it into the turbine part.

Another object of the invention is to introduce cooling air into the rotating turbine section without creating excessive tangential flow velocities.

In one embodiment of the invention, air diverting shaping vanes are mounted to the shaft wall in correspondence with holes in the shaft wall. The purpose of these vanes is to divert the direction of flow of the cooling air, which is accelerated from the nozzle in a direction tangential to the shaft wall, to a new direction parallel to the centerline of the shaft hole. As a result, the vanes serve to introduce cooling air into the holes of the rotating shaft without causing large pressure losses and with a lower resulting cooling air temperature.

Additionally, the passages between the vanes are enlarged in the direction of airflow to diffuse air and increase static pressure. Reduce inlet losses by providing an aerodynamically shaped inlet. This reduction in inlet losses, combined with pressure recovery by diffusion, allows the use of higher pressure ratio nozzles, thereby increasing the efficiency of the system and lowering the exit temperature of the cooling air to the rotor. The reduction in cooling air temperature allows the cooling air flow to be reduced, ultimately increasing turbine cycle efficiency.

Next, the present invention will be explained in detail with reference to the drawings.

A typical gas turbine aircraft engine 10 is shown in FIG. 1 for the purpose of illustrating the basic components and function of the engine, as well as providing a general description of the cooling air flow path. In basic engine function, incoming air first enters the compressor 12, which compresses the air to extremely high pressures that support rapid combustion downstream of the engine. This highly compressed air is directed via compressor outlet 14 to combustor 16 where the air is mixed with fuel and ignited. The ignited air-fuel mixture forms high-temperature combustion gas, and the combustion gas is transferred to combustor 1.
6 and accelerates into the turbine section 18. In the turbine section, accelerated combustion gases are directed to turbine blades 20 causing the blades to rotate at extremely high speeds. Turbine blades 20 are coupled to a turbine shaft 24 along with a rotor 22 and configured to transmit power to the turbine shaft. Shaft 24 can be mechanically coupled to any device that the engine user desires to mechanically drive. A typical aircraft engine uses a turbine shaft to drive the compressor 12 as well as a fan (not shown) that accelerates the air to provide forward thrust to the aircraft.

When performing these basic engine functions, it is important to understand that maximum power can be extracted from the combustion gases at a specific thermodynamically determined optimum temperature. Unfortunately, the calculated optimum temperature is too high, and when the engine is operated in its optimum mode,
Engine parts exposed to hot combustion gases are quickly destroyed. Therefore, gas turbine engines actually operate at temperatures somewhat lower than the thermodynamically determined optimum level.

In an effort to improve engine efficiency by enabling higher temperatures, recent design ideas have focused on improving air-cooled turbine components within the combustion flow path. These attempts have been successful, resulting in significant increases in the efficiency of modern gas turbine engines. However, this cooling air must be extracted from a high pressure source, such as the engine compressor 12, and the air extracted from the engine compressor results in parasitic losses of air available for combustion.
This therefore leads to a loss of engine power output.

In view of this parasitic loss, much effort has been devoted to methods of handling this cooling air so that less air must be removed from the compressor 12 to cool the hot turbine components.

FIG. 2 shows a portion of a conventional engine and its internal cooling air passages. The path of the cooling air flow is indicated by a thick, blurred arrow. Cooling air is supplied to the compressor outlet (second
(not shown) into the region 25 surrounding the combustor wall 26 and finally through the turbine shaft 24 and into the turbine rotor cavity 19 . This section is a critical area in the cooling air flow path as the air is directed to the hottest turbine components, including the first row of turbine blades 30. Air must be maintained at high pressure. This is because in order for the air to flow into and past the blades 20, the air must be at a higher pressure than the combustion gases surrounding the nozzle and blades. Because the combustion gases have just exited the compressor 12 and combustor 16, the combustion gases are at a relatively high pressure relative to the rest of the engine.

Although this part is a critical area of the cooling air flow path, the technical difficulties of moving air in this area of the engine are more complex. This is because air flows from the non-rotating parts of the engine through the rotating shaft 24 and into the rotor cavity 19. As the air enters the turbine shaft 24 through the plurality of holes 32 in the wall of the turbine shaft 24, it must be rapidly accelerated in the direction of rotation. Where the non-rotating air passes through the holes 32,
Significant inefficiencies and pressure changes occur in this region.

In an attempt to reduce such inefficiencies, engine designers have recently incorporated nozzles 34 into cooling air passages of the type shown in FIG. Nozzle 3
4 to accelerate the cooling air in the direction of rotation of the turbine. As a result of this acceleration, the air is forced to flow in a direction tangential to the circumference of the shaft. If the air has a tangential flow velocity greater than the tangential velocity of the shaft, large pressure losses will occur as the air passes through the holes in the shaft. Additionally, as when air passes under the first stage turbine rotor 22,
If it flows into a smaller radius, its tangential direction increases and this causes acoustic resonance. Eliminating excess tangential velocity before the shaft hole significantly improves this situation. This was cited as an objective of the present invention.

FIG. 3 shows one embodiment of the present invention. The device of the invention is incorporated in the area of the engine 10 where the cooling air enters the turbine shaft 24. According to the invention, where air enters the shaft 24,
Deswirler in the air flow path (deswirler de-swirl device)
Attach 36. The function of the deswirler 36 is to aerodynamically redirect the cooling air to guide the air into the holes 32. The de-swirler 36 also functions to reduce the rotational speed of the cooling air so that the magnitude of the rotational speed of the cooling air approaches the rotational speed of the turbine shaft 24 . death swirler 3
6 is attached directly to the turbine shaft 24, so
It rotates in the same manner as the shaft 24. This feature allows the de-swirler 36 to reduce the rotational speed of the cooling air as it directs the air into the holes 34.

As seen in FIG. 3, another feature of the de-swirler 36 of the present invention is that the multiple passageways 38 of the de-swirler 36 are formed in a cross-section that widens from the inlet 40 to the outlet 42. These Suehiro passages 3
8 acts as a diffuser, thereby converting a portion of the air approach velocity head to static pressure. Again, it must be emphasized that this cooling air is introduced into the high pressure area of the turbine. It is highly desirable to maintain the cooling air at a high pressure at the point where the air enters the hole 32 in the shaft. Therefore, in the particular application of the de-swirler 36 shown in FIG. 3, it is highly desirable to have its internal passageway 38 diverging to provide this diffusion function.

In FIG. 4, the nozzle 34, de-swirler 36 and shaft hole 32 are illustrated so that the directing or guiding effect of the nozzle and de-swirler on the cooling air flow is readily understood. The cooling air flow path is indicated by a thick, blurred arrow. Following the course of this flow path, the nozzle 34 is aligned in a substantially serial flow relationship with the deswirler inlet 40, and the deswirler outlet 42 is also aligned in a similar relationship with the shaft hole 32.
This provides a substantially aerodynamic flow path for cooling air.

As previously mentioned, the cooling medium must be drawn from a relatively high pressure air source within the engine.
One ideal location is the region 25 surrounding the combustor wall just downstream of the compressor outlet. This air is at very high pressure and its extraction location is immediately upstream of the turbine section 18, allowing passage to the turbine shaft 24.

The first step in transferring this air to the turbine section 18 is to accelerate the air in the direction of rotation of the turbine. As previously mentioned, this is accomplished with nozzle 34. The operation of nozzles is well known to those skilled in the art, and any of a variety of types can be used for air acceleration. The degree of acceleration of the cooling air can be controlled by changing the nozzle structure.

After passing through the nozzle outlet 44, the air is directed to the de-swirler inlet 40. The de-swirler 36 consists of a series of diverting vanes 37 that define a passageway 38 that diverts the air from a tangential direction to a direction more parallel to the centerline of the hole 32.

The vanes 37 accomplish this redirection by simultaneously bending the airflow radially inward and changing a portion of the tangential velocity of the airflow to a rotational speed that generally matches the rotational speed of the turbine shaft 24. achieve the goal.

Moreover, if the tangential velocity of the air significantly exceeds the rotational speed of the turbine shaft, the passage 3 of the deswirler is configured to diffuse the air.
8 functions to convert a portion of the approaching flow head to static pressure and to reduce pressure loss upon shaft entry. This low entry loss due to diffusion and pressure recovery combine to provide a higher pressure ratio and acceleration across the nozzle 34. The higher the pressure ratio across the nozzle 34, the lower the air temperature at the nozzle outlet 44. The reduction in cooling air temperature allows the cooling air flow to be reduced, thus increasing turbine efficiency. This is the main objective of the invention. A reduction in cooling air temperature is achieved by transferring a portion of the air's energy to the turbine, further increasing turbine efficiency.

[Brief explanation of drawings]

Fig. 1 is a schematic cross-sectional view of a typical gas turbine aircraft engine, Fig. 2 is a cross-sectional view of a portion of the gas turbine engine showing a typical conventional cooling air passage, and Fig. 3 is a schematic cross-sectional view of a typical gas turbine aircraft engine. 2, and FIG. 4 is a cross-sectional view taken along line 4--4 in FIG. 3. 10... Gas turbine engine, 18... Turbine part, 20... Turbine blade, 22...
Rotor, 24... Turbine shaft, 32...
Hole, 34... Nozzle, 36... Death swirler, 37
... Vane, 38 ... Passage, 40 ... Entrance, 42
...Exit, 44...Nozzle exit.

Claims (1)

Claims: 1. A compressor, a combustor, a rotating turbine section located about a central engine axis and having a rotating turbine shaft and a turbine rotor, and a cooling air system for delivering cooling air to the rotating turbine section. , wherein the cooling air system has a nozzle that accelerates cooling air tangentially to the rotating turbine shaft and in the direction of rotation of the turbine, the cooling air system receiving the accelerated cooling air from the nozzle, means for introducing cooling air into the bore of the rotating turbine shaft without significant cooling air pressure losses; A cooling air guide device for a turbine, comprising a rotating deswirler having a plurality of curved passages, said curved passages having a gradually expanding cross section to diffuse air and convert a portion of the approaching flow velocity head of the air into static pressure. . 2. The apparatus of claim 1, wherein the outlet of the passageway is in series flow relationship with the centerline of the receiving hole of the rotating turbine shaft. 3. The device according to claim 1 or 2, wherein the deswirler is directly attached to the rotating turbine shaft and rotates together with the shaft. 4. A compressor, a combustor, a rotating turbine section located about a central engine axis and having a rotating turbine shaft and a turbine rotor, and a cooling air system for delivering cooling air to the rotating turbine section, the cooling air system comprising: A gas turbine engine having a nozzle for accelerating cooling air tangentially to the rotating turbine shaft and in the direction of rotation of the turbine, comprising a diffuser type deswirler that is directly attached to the turbine shaft and rotates together with the shaft, the deswirler a plurality of internal vanes forming a plurality of passages having progressively curved inner surfaces, the passages comprising: (a) said curved inner surfaces aligned in series flow relationship with an accelerated cooling air flow exiting said nozzle to aerodynamically direct said cooling air; (b) an intermediate region having a progressively increasing cross-sectional area to slow and diffuse the cooling air; (c) the curved inner surface opening into the turbine shaft; and an outlet region between the vanes configured to aerodynamically introduce the cooling air into the turbine shaft aligned in a series flow relationship with accelerated cooling air without significant pressure loss. A device that introduces water into the rotating turbine section.