JPS6236142B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to gas turbine engines, and more particularly to temperature conditioning of a hot gas stream exiting a combustor of a gas turbine engine.

Current gas turbine engines used as aircraft prime movers operate at high gas temperatures. In fact, one of the main performance factors that determines the thrust of an engine is the combustor exit temperature. In order to generate and maintain a certain rated thrust, there must be a hot or average gas temperature exiting the combustor, which temperature level is typically the highest average gas temperature encountered within the engine. In many cases, this temperature level approaches the critical temperature of components such as turbine vanes located at the combustor outlet. As a result, designers are faced with the problem of balancing turbine vanes with high average temperatures of the hot gases exiting the combustor.

The non-uniformity of the hot gas temperature at the combustor exit surface is
This is another factor that makes the above-mentioned problem of coexistence even more difficult for designers. Temperature non-uniformity is generally due to the geometrical design of the combustor itself. For example, combustor fuel injectors generate localized temperatures at the exit surface that are significantly higher than the average temperature, contributing to non-uniform exit temperature distribution. In particular, during the combustion process, combustion of the air-fuel mixture tends to be relatively intense at the fuel injection point within the combustor. Because the airflow through the combustor is high velocity, these intense combustion zones extend into hot bands that extend axially along the length of the combustor.
In many cases, the hot band can extend axially aft enough to encompass turbine vanes downstream of the combustor outlet. Therefore, while it is generally true that designers must design for the average temperature of the hot gas at the exit face of the combustor, designers effectively design turbine vanes for the highest temperature of the hot gas at the exit face. It is necessary to design it so that it is compatible with the point temperature. Therefore, this single point temperature is a significant problem for designers who have typically addressed it by applying existing cooling techniques. In particular, standard strategies have been, for example, film cooling of the surfaces of the vanes, or impingement cooling and internal convection cooling of the vanes using compressor exhaust air. However, when cooling air is used in this manner, engine performance is degraded in the form of reduced thrust or increased fuel consumption per unit thrust. Furthermore, in these conventional devices, the cooling air is introduced at a location where the gas exhibits a high Matzha number, resulting in large mixing losses. Additionally, vanes designed using cooling techniques such as impingement cooling are complex and expensive components in modern turbine engines.

The degree of non-uniformity of the temperature distribution of the gas at the exit face of the combustor is highly dependent on the length allowed for combustion. Short combustors are prone to relatively high strip temperatures due to improper mixing length. Accordingly, some conventional engines utilize relatively long combustors to eliminate the effects of uneven temperature distribution. The present invention provides a preselected circumferential distribution of hot gases at the combustor exit surface,
This distribution reduces the required flow rate of vane cooling air, simplifies the mechanical construction of the vanes, and allows for shorter combustor designs.

Accordingly, it is an object of the present invention to make turbine vanes of a gas turbine engine compatible with hot gases exiting the engine combustor.

Another object of the invention is to ensure compatibility between turbine vanes and hot gases exiting the fuel without adversely affecting engine performance.

Another object of the present invention is to eliminate the adverse effects of hot gases exiting an engine combustor on uneven temperature distribution turbine vanes.

Another object of the invention is to provide the hot gases exiting the relatively short combustor with a temperature distribution that is pre-qualified to be compatible with turbine vanes.

Briefly stated, to accomplish the above and other objects of the invention which should be apparent from the following description and accompanying drawings, the invention in one aspect comprises To provide a gas turbine engine in which hot gas flows in an annular passage defined by a gas turbine engine having the following improved structure. That is, means for establishing a preselected circumferential temperature gradient in the hot gases are provided upstream of the turbine vanes associated with the combustor. The temperature gradient is preselected so that a relatively high temperature hot gas flow passes through the gap between the stator vanes and a relatively low temperature hot gas flow impinges on the stator vanes. This gradient may be established by providing a first group of holes and a second group of holes as a means of introducing air into the engine's combustor. The first group of holes has a larger cross-sectional area than the second group of holes and is disposed in axial alignment with the turbine stator blade, and the second group of holes is disposed in axial alignment with the gap. . The sinusoidal waveform characteristics of the present invention are enhanced by using a multiple of the number of stator vanes as the number of fuel injectors delivering fuel to the engine.

The invention will now be described with reference to the accompanying drawings.

FIG. 1 depicts a typical air-breathing gas turbine engine with a total of 30
FIG. The engine 30 has an inlet 32
, a booster assembly 34 , and a compressor assembly 36
, a combustor assembly 38, a turbine assembly 40, and an exhaust section 44, these components being arranged in series. Internal annular channel 3 extending in the axial direction
3 extends rearwardly from inlet 32 to exhaust 44 and serves as a flow path for air flowing through engine 30 .
Also, a plurality of turbine vanes 46 forming part of the turbine assembly 40 are disposed within the annular flow passage 33 immediately downstream of the combustor assembly 38 .
Ambient air entering inlet 32 is compressed by booster 34 and compressor 36. This compressed air enters the combustor 38 where it is mixed with fuel and combustion occurs. The fuel gas produced by this combustion is 2,500〓 in some types of gas turbine engines.
(1370℃). After exiting the combustor 38, the combustion gases pass through turbine vanes 46 and through the remainder of the turbine assembly 40. Turbine assembly 40 extracts energy from the hot combustion gases to drive booster 34 and compressor 36.
The hot gases then exit the engine 30 through the exhaust 44 at high velocity, whereby the residual energy of the hot gases assists the engine 30 in generating thrust.

FIG. 2 is a perspective view of combustor assembly 38 in functional association with a plurality of turbine vanes 46 and the remainder of turbine assembly 40. FIG. Combustor assembly 38
includes an axially and circumferentially extending outer liner assembly 48 and an inner liner assembly 50, the flow liner assemblies being radially spaced apart and mutually defining a portion of the annular flow passage 33. combustor liner 48,
At the upstream end of 50 , a plurality of fuel injectors 52 are mounted within a plurality of holes 54 in combustor assembly 38 .
Note that the combustor assembly 38 is preferably annular and extends circumferentially about the engine centerline. The fuel injectors 52 thus provide a plurality of injection points that are circumferentially spaced and direct the air/fuel mixture into the combustor assembly 38 along the entire circumference of the annular passageway 33 .

Outer liner 48 is comprised of a plurality of integrally formed stepped hoops 56. Each hoop has a generally axially extending generally cylindrical portion 58 and a radially and circumferentially extending integral step 60 disposed at its downstream end. The step 60 is integrally coupled to the upstream end of the next adjacent downstream hoop 56.
The lip 62 at the downstream end of each upstream hoop 56 partially overlaps the upstream end of the next adjacent downstream hoop 56, thus serving as a means for film cooling the inner surface of the liner 48.

Similarly, inner liner 50 is comprised of a plurality of integrally formed stepped hoops 64. Each hoop has a generally axially extending cylindrical portion 66 and a radially and circumferentially extending integral step 68 disposed at its downstream end. The step 68 is integrally connected to the upstream end of the next adjacent downstream hoop 65. The lip 70 at the downstream end of each upstream hoop 64 partially overlaps the upstream end of the next adjacent downstream hoop 64, thus serving as a means for film cooling the interior surface of the liner 50.

A plurality of turbine vanes or nozzle vanes 64 are positioned immediately downstream of the combustor assembly 38 and are spaced apart along the entire circumference of the combustor assembly 38 . The vanes 46 are secured to an outer vane platform or shroud 72 and an inner vane platform or shroud 73, both of which further define the aforementioned flow path 33. Immediately downstream of the stator vanes, a plurality of circumferentially spaced turbine rotor blades 74 are mounted on a rotatable rotor disk 76 . The rotor blades 74 receive energy from the hot gas flow passing through the stator blades 46 .

As air introduction means, a first group of dilution holes 78 and a second group of dilution holes 80 that are spaced apart in the circumferential direction are provided in each of the outer liner 48 and the inner liner 50.
Each hole 80 is interposed between two adjacent holes 78 and has a smaller cross-sectional area than each hole 78 . Both dilution holes 78 and 80 introduce additional air into the combustor 38. This additional air mixes with the air/fuel mixture from injector 52 to accelerate and complete the combustion process.

Next, referring to FIG. 3, the circumferential positions of the dilution holes 78 and 80 determined for the turbine stator blade 46 according to the present invention will be described. FIG. 3 is a diagram illustrating the axial alignment of dilution holes 78, 80 with respect to turbine vane 46, which alignment constitutes a part of the present invention. The horizontal axis of this figure schematically indicates a circumferential position around the centerline of the gas turbine engine, and the vertical axis schematically indicates an axial position along the centerline of the gas turbine engine. Note that the position moves from the front to the rear as you move from the bottom to the top of this diagram. Thus, as can be seen, the leading edges 82 of the turbine vanes 46 are spaced approximately 40 degrees apart along the entire circumference of the engine. Also, as shown, the dilution hole 78 is disposed in axial alignment with the leading edge 82 of the turbine vane 46 . For example, at a circumferential position of 120 degrees, one hole 78 is axially aligned with and immediately upstream of the leading edge of one vane 46. In either case, there is a dilution hole 78 just upstream of the leading edge 82 of the vane 46 .

On the other hand, the dilution hole 80 is located in the gap 8 between adjacent leading edges 8.
4 in axial alignment. for example,
One dilution hole 80 connects the leading edge 82 at a 120° position.
It is axially aligned with a gap 84 between the leading edge 82 at a 160° position. This alignment is repeated at various circumferential locations about the centerline of the engine.

The average turbine inlet temperature 88 is also shown schematically in FIG. 3 together with a continuous circumferential temperature distribution curve 90. Curve 90 represents the temperature at various circumferential separations of the inlet face to vane 46 . As shown, in accordance with the present invention, the holes 78, 80 are axially aligned with the leading edge 82 and gap 84, respectively, so that the temperature of the hot gas impinging the vane is lower than the average gas temperature 88; In contrast, the temperature of the hot gas flowing axially into the gap 84 is higher than the average gas temperature 88. That is, a preselected circumferential temperature gradient is established such that relatively high temperature hot gas flow passes through gap 84 and relatively low temperature hot gas impinges on vane 46. This sinusoidal characteristic of the continuous temperature distribution curve is based on the fact that the dilution air entering the combustor has a lower temperature than the air-fuel mixture during combustion and therefore has cooling capacity. Large holes 78 allow a relatively large amount of air to pass through where they are axially aligned with vanes 46, whereas small holes 80 allow a relatively small amount of air to pass through where they are axially aligned with gap 84. The large holes 78 provide a greater cooling effect than the small holes 80. Thus, the larger cross-sectional area dilution holes 78 create localized cold spots or cold bands. These cold spots or stripes are located at the leading edge 82.
The cooling-enhancing properties of the dilution temperature can be utilized by matching the dilution temperature. To explain in more detail, dilution hole 7
By arranging 8 and 80 as described above,
A suitable circumferential temperature distribution or gradient with alternating high and low temperature regions can occur. The hot region corresponds to the interblade gap 84, and the stator blades 46 themselves are located within the cold region. Because the holes 78, 80 are arranged as described above, the average turbine inlet temperature, which is a measure of the engine's available thrust, is maintained at the desired level. This is because the high temperature gas flow between the vanes 46 is supplemented by a relatively low temperature gas flow. Therefore, the specific arrangement of the dilution holes described above allows the temperature of the hot gas at the leading edge of the vane 46 to be reduced without affecting the average temperature of the hot gas at the combustor exit annulus.

Another advantage of the above-described arrangement resides in the fact that the localized high temperature or hot streak created by the injector 52 cannot penetrate the low temperature zone established in front of the leading edge 82, as described above. Additionally, by using a multiple of the number of vanes 46 as fuel injectors 52, the sinusoidal nature of continuous temperature curve 90 is reinforced. Such a configuration facilitates temperature domain control by allowing a particular formulation sequence to be repeated around the circumference of the engine.

The device described above is suitable for achieving the above object of the invention. The present invention eliminates the randomness of temperature gradients that existed in conventional devices.
That is, an ordered or preselected temperature gradient is established instead of the conventional random temperature gradient, and the hot portion of the gas flow is confined to the gap between the vanes. This is achieved by controlling the existing dilution air to promote the combustion process.
The present invention does not require additional cooling air and elaborate stator vane cooling designs as required by conventional devices. The invention therefore has considerable advantages over conventional devices known to those skilled in the art.

[Brief explanation of the drawing]

FIG. 1 is a schematic diagram of a typical gas turbine engine to which the present invention is applied, FIG. 2 is a partially enlarged perspective view of the combustor section and turbine section of the engine shown in FIG. 1 is a graph showing the relative axial position of elements according to the invention; 46... Turbine stator blade, 48, 50... Outer and inner liners, respectively, 52... Fuel injector,
72, 73...outer and inner stator blade shrouds, respectively, 78, 80...dilution hole, 84...blade gap, 88...average turbine inlet temperature, 90...
...Circumferential temperature distribution curve.

Claims (1)

Claims: 1. Hot gas flows within passageways defined in part by inner and outer liners and inner and outer stator vane shrouds of a combustor, in which air and fuel flow. in a gas turbine engine having a plurality of injectors at the upstream end of the combustor for introducing the raw mixture into the combustor to combust the raw mixture, the hot gas having an average turbine inlet temperature. multiples of the number of circumferentially spaced turbine stator vanes are disposed in the hot gas flow path downstream of the combustor, each stator vane extending radially across the hot gas flow path; The spacing between the stator vanes forms an inter-stator vane gap, a first plurality of circumferentially spaced holes are disposed in axial alignment with the stator vanes, and a second plurality of circumferentially spaced holes are disposed in axial alignment with the stator vanes. spaced apart holes are disposed in axial alignment with the vane gap, the holes for introducing additional air into the combustor to mix with the raw mixture and promote combustion. disposed in the combustor, and the holes are positioned such that combustion of the additional air and raw mixture creates a preselected temperature gradient in the hot gas, the temperature gradient being A gas turbine engine characterized in that the hot gas flowing through has a first temperature above the turbine inlet average temperature and the hot gas impinging on the stator blade has a second temperature below the turbine inlet average temperature. . 2. The gas turbine engine according to claim 1, wherein the cross-sectional area of one of the second plurality of holes is smaller than the cross-sectional area of one of the first plurality of holes.