Patent Description:
The present disclosure relates generally to turbine engines and, more specifically, to combustor architectures that provide desired downstream conditions.

A conventional gas turbine engine typically includes a compressor for compressing air that is mixed with fuel and ignited in a combustor for generating a high pressure, high temperature gas stream, referred to as combustion gas. The combustion gases flow to a turbine, where they are expanded, converting thermal energy from the combustion gases to mechanical energy for driving a shaft to power the compressor and produce output power for powering an electrical generator or to produce thrust in Aviation applications, for example.

In at least some known gas turbines, a first set of guide vanes (or diffuser) is coupled between an outlet of the compressor and an inlet of the combustor. The first set of guide vanes reduces angular momentum, thus reducing swirl (i.e., removing bulk swirl) and flow angle of a flow of air discharged from the compressor such that the flow of air is channeled in a substantially axial direction towards the combustor. A second set of guide vanes (or first stage turbine nozzle) is coupled between an outlet of the combustor and an inlet of the turbine. The second set of guide vanes facilitates increasing angular momentum, swirl (i.e., reintroducing bulk swirl) of a flow of combustion gas discharged from the combustor such that flow angle requirements for the inlet of the turbine are satisfied. However, redirecting the flows of air and combustion gas with the first and second sets of guide vanes increases operating inefficiencies of the gas turbine. Moreover, including additional components, such as the first and second sets of guide vanes generally adds weight, cost, and complexity to the gas turbine.

Bulk swirl combustors (opened or confined) may be used to provide the necessary swirl, flow angles and flow conditions (i.e., angular momentum) thereby reducing or eliminating the need for either the first or second sets of guide vanes, or possibly both.

<CIT> describes a turbine engine. <CIT> describes a gas turbine combustion chamber. <CIT> describes an internal-combustion turbine power plant. <CIT> describes a combustion chamber for a gas turbine.

According to the invention, there is provided a gas turbine engine as claimed in claim <NUM>.

Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments.

Here and throughout the specification and claims, range limitations may be combined and/or interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.

As used herein, the term "axial" refers to a direction aligned with a central axis or shaft of the gas turbine engine. An axially forward end of the gas turbine engine is the end proximate the fan and/or compressor inlet where air enters the gas turbine engine. An axially aft end of the gas turbine engine is the end of the gas turbine proximate the engine exhaust where low pressure combustion gases exit the engine via the low pressure (LP) turbine.

As used herein, the term "circumferential" refers to a direction or directions around (and tangential to) the circumference of an annulus of the gas turbine, or for example the circle defined by the swept area of the turbine blades. As used herein, the terms "circumferential" and "tangential" are synonymous.

As used herein, the term "radial" refers to a direction moving outwardly away from the central axis of the gas turbine. A "radially inward" direction is aligned toward the central axis moving toward decreasing radii. A "radially outward" direction is aligned away from the central axis moving toward increasing radii.

As used herein, the term "bulk swirl" refers to swirling of compressed air and/or combustion gases within a gas turbine annulus around an engine centerline.

As used herein, the term "confined swirl" refers to structures used to increase the magnitude of bulk swirl (angular momentum) induced by heat addition in the combustor section.

As used herein, the terms "angular momentum," swirl," "tangential flow," and "flow angle" describe engineering principles that are correlated.

<FIG> is a schematic illustration of an exemplary turbine engine <NUM> including a fan assembly <NUM>, a low-pressure or booster compressor assembly <NUM>, a high-pressure compressor assembly <NUM>, and a combustor assembly <NUM>. The combustor assembly <NUM> includes a contoured flow-path 18A of the present embodiments and an exemplary gas turbine engine combustor flow-path 18B (illustrated with dashed-lines in <FIG>). The magnitude of the contouring of the contoured flow-path 18A of the present embodiments is dramatically increased in contrast to exemplary gas turbine engine combustor flow-path 18B. Fan assembly <NUM>, booster compressor assembly <NUM>, high-pressure compressor assembly <NUM>, and combustor assembly <NUM> are coupled in flow communication. Turbine engine <NUM> also includes a high-pressure turbine assembly <NUM> coupled in flow communication with combustor assembly <NUM> and a low-pressure turbine assembly <NUM>. Turbine engine <NUM> has an intake <NUM> and an exhaust <NUM>. Turbine engine <NUM> further includes a centerline <NUM> about which fan assembly <NUM>, booster compressor assembly <NUM>, high-pressure compressor assembly <NUM>, and turbine assemblies <NUM> and <NUM> rotate.

In operation, air entering turbine engine <NUM> through intake <NUM> is channeled through fan assembly <NUM> towards booster compressor assembly <NUM>. Compressed air is discharged from booster compressor assembly <NUM> towards high-pressure compressor assembly <NUM>. Highly compressed air is channeled from high-pressure compressor assembly <NUM> towards combustor assembly <NUM>, mixed with fuel, and the mixture is combusted within combustor assembly <NUM>. High temperature combustion gas generated by combustor assembly <NUM> is channeled towards turbine assemblies <NUM> and <NUM>. Combustion gas is subsequently discharged from turbine engine <NUM> via exhaust <NUM>. A high-pressure shaft <NUM>, is concentrically disposed about the centerline <NUM> and mechanically couples the high-pressure compressor assembly <NUM> to the high-pressure turbine assembly <NUM>. The embodiments, combustor and turbine engine described herein are applicable to several possible engine architectures including, but not limited to, turboshaft engines, aircraft engines, turboprop engines, turbofan engines, turbojet engines, geared architecture engines, direct drive engines, land-based gas turbine engines, etc..

<FIG> illustrates an embodiment of a gas turbine engine 10A including can-style combustor assemblies 18B, an engine architecture to which the present embodiments are also applicable.

<FIG> is a diagrammatic schematic of the axially aft end <NUM> of the compressor, the combustor and the axially fore end <NUM> of the turbine engine <NUM>. The compressed airflow exists the last stage compressor rotor <NUM> at compressor rotor exit angle <NUM>, αA and enters the compressor exit guide vane <NUM> where the flow is turned in an axial direction <NUM> at station B (angular momentum decreases). Between station Band C, heat <NUM> is added in the combustor section <NUM>. At station C, the combustor flow exits the combustor in an axial direction <NUM> and enters the first stage turbine nozzle <NUM> where the flow is redirected to a turbine inlet angle <NUM>, αD (angular momentum increases). Stated otherwise, the flow enters the compressor exit guide vane <NUM> where angular momentum decreases only to then be increased at the first stage turbine nozzle <NUM>.

<FIG> is a diagrammatic schematic of the axially aft end <NUM> of the compressor, the combustor and the axially fore end <NUM> of the turbine without a last stage compressor vane and without a first stage turbine nozzle. The airflow exits the last stage compressor rotor <NUM> at compressor rotor exit angle <NUM>, αA and enters the combustor section <NUM> where heat <NUM> is added. The combustor flow exits the combustor <NUM> and enters the first stage turbine rotor <NUM> at turbine inlet angle <NUM>, αD. In operation, the flow angle may change as heat <NUM> is added in the combustor <NUM> to the compressed air in the form of ignited fuel. The change in flow angle may differ at various operating conditions from low power to high power, but generally the flow angle will decrease with heat addition. Therefore, it may be desirable for flow to exit the last stage compressor rotor <NUM> with increased flow angle (αA > αD) such that flow enters the first stage turbine rotor <NUM> at the correct angle and flow conditions.

<FIG> is a cross-section side view of an exemplary combustor section <NUM> of the present embodiments, including the axially aft end <NUM> of the compressor <NUM>, and the axially forward end <NUM> of the turbine section <NUM>. The combustor section <NUM> of <FIG> may be used in connection with the embodiment of <FIG> in which neither a last stage compressor vane <NUM> nor a first stage turbine vane (or nozzle) <NUM> is required to provide the required flow angles. The embodiment of <FIG> provides a combustor geometry that balances the flow angle, among other parameters (e.g., angular momentum, temperature and pressure). The combustor geometry is comprised of three sections (a combustor inlet <NUM>, a combustor heat addition portion <NUM>, and combustor exit <NUM>) with purposefully designed flow path contouring such that conservation of angular flow momentum and flow changes with heat addition will yield the required conditions at the turbine rotor inlet, while minimizing losses and cooling requirements. The flow path contouring varies with axial direction in such a manner as to purposefully achieve specific flow conditions at various stations throughout the combustor section <NUM>, via a compressor mean radius 74B, a combustor mean radius <NUM>, a turbine mean radius <NUM> and span <NUM>.

The combustor inlet <NUM> curves radially outward to reduce the flow Mach number. The combustor heat addition portion <NUM> further adapts the span <NUM> to further manage Mach numbers and hence minimize losses (Rayleigh) and cooling requirements. Stated otherwise, the span is another design parameter that can be varied to achieve the appropriate flow conditions at each station. In one embodiment, the gas turbine engine <NUM> may include an intentionally-varied span <NUM> to achieve the appropriate flow conditions at each station (minimizing losses and providing the appropriate flow conditions to the first stage turbine rotor <NUM>). Thus, the present embodiments may enable the elimination of the first stage nozzle, due to the radius variation and/or span tailoring. The combustor exit <NUM> curves radially inward to create the appropriate flow angle (swirl) into the first stage turbine rotor <NUM>. This out and back-in configuration would be desired even in examples where the compressor exit at station A and first stage turbine inlet <NUM> at station D are at roughly the same radii (i.e., compressor mean radius 74B may be approximately equal to turbine mean radius <NUM> (for example within about <NUM>% or about <NUM>%) or may be different than the turbine mean or mid-span radius <NUM>). This differs from the state-of-the-art in which the combustor roughly connects the compressor exit to the turbine inlet in a direct and substantially monotonic fashion (i.e., continually increasing or decreasing radius).

As illustrated in <FIG>, the combustor section <NUM> stretches from station A to station D and includes a combustor inlet flow path <NUM> that curves radially outward from an exit of the last stage compressor rotor <NUM>. Both a combustor inlet inner wall 52A and the combustor inlet outer wall 52B curve steeply in a radially outward direction <NUM> within the combustor section <NUM>. The combustor inlet inner wall 52A and the combustor inlet outer wall 52B define the radially inner and radially outer boundaries of the combustor inlet flow path <NUM>, respectively. Upon reaching the heat addition portion <NUM>, a combustor inner wall 54A and a combustor outer wall 54B define a combustor cavity <NUM> in which heat <NUM> is added in the form of ignited fuel.

Still referring to <FIG>, the combustor inner wall 54A and the combustor outer wall 54B are formed such that the annular combustor cavity <NUM> defines a combustor heat addition portion <NUM> which includes a substantially axial orientation. In addition, the combustor heat addition portion <NUM> has a combustor mean radius <NUM> which is the average of the radius of the combustor inner wall 54A and the combustor outer wall 54B, defined at the combustor heat addition portion <NUM>. The radial distance between the combustor inner wall 54A and the combustor outer wall 54B defines a combustor span <NUM>. The combustor section <NUM> includes a combustor exit <NUM> downstream of the combustor inner wall 54A and the combustor outer wall 54B such that combustion gases flow into the combustor exit <NUM> in which both a combustor exit inner wall 66A and a combustor exit outer wall 66B define a combustor exit flow path <NUM>, which curves radially inward.

Referring still to <FIG>, the combustor flow exits the combustor exit <NUM> at station D and enters the turbine section <NUM> at the first stage turbine blade <NUM>. A first stage turbine flow path <NUM> is defined between a turbine outer wall 72B and a turbine inner wall 72A, the turbine outer wall 72B defining the radially outer boundary of the turbine flow path <NUM> and the turbine inner wall 72A defined the radially inner boundary of the turbine flow path <NUM>. A turbine mid-span radius <NUM> is defined as the average of the radius of the turbine inner wall 72A and the radius of the turbine outer wall 72B, both defined at a first stage turbine rotor (or blade) leading edge <NUM>. In one embodiment, the combustor mean radius <NUM> is between about <NUM> and about five times the turbine mid-span radius <NUM>. In another embodiment, the combustor mean radius <NUM> may be between about <NUM> and about <NUM> times the turbine mid-span radius <NUM>. In another embodiment, the combustor mean radius <NUM> may be between about <NUM> and about <NUM> times the turbine mid-span radius <NUM>. As such, the combustor <NUM> of the present embodiments is a large radius combustor. In another embodiment, the gas turbine includes inner and outer flow paths that are curved radially outwardly upstream of the heat addition in the combustor section and radially inwardly downstream of heat addition in the combustor section. In another embodiment, the combustor mean radius <NUM> is greater than both the turbine mid-span radius <NUM> and compressor mean radius 74B. In another embodiment, the combustor mean radius <NUM>, the turbine mid-span radius <NUM>, and compressor mean radius 74B are all different (i.e., they all include different radii). The gas turbine inner and outer flow paths include the combustor inlet inner and outer flow paths, 52A and 52B, the combustor inner and outer flow paths, 54A and 54B, and the combustor exit inner and outer walls, 66A and 66B, respectively.

<FIG> is a diagrammatic schematic of the axially aft end <NUM> of the compressor section <NUM>, the combustor section <NUM>, and the axially fore end <NUM> of the turbine section <NUM> with a last stage variable compressor vane (VCV) <NUM> and without a first stage turbine nozzle. Including a last stage variable compressor vane (VCV) <NUM> enables the flow angle and angular momentum to be selectively controlled to account for varying operation conditions of the gas turbine engine <NUM>. For example, the geometry of the combustor section <NUM> as illustrated in <FIG> may be contoured so as to provide a desired flow angle at the inlet of the first stage turbine rotor <NUM> at a specific design point. However, when the operation of the gas turbine engine deviates from the specific design point, variable compressor vanes <NUM> may be modulated to produce the desired flow angle at the first stage turbine rotor <NUM>. Additionally, these variable vanes may be used to change the flow rate through the machine. Thus, moving the standard choke point of the machine from aft of to before the heat addition point. Stated otherwise, the variable compressor vane <NUM> modulates the VCV flow area such that the VCV flow area is less than a minimum flow area of the turbine section. This may result in gaining the benefits of a variable area turbine (VAT) or variable area turbine nozzle (VATN) without the challenges associated with variable geometry components in the hot section. In the embodiment of <FIG>, air exits the compressor section <NUM> after passing through the last stage variable compressor vane (VCV) <NUM> and enters the combustor section, where heat is added to the air in a combustion process. Combustion gases then exit the combustor section <NUM>, and enter the turbine section <NUM> directly at the first stage turbine blade <NUM> without first passing through a stage <NUM> nozzle. Stated otherwise, the first stage turbine blade <NUM> is upstream of every nozzle stage within the turbine section <NUM>. Similarly, every compressor rotor stage disposed within the compressor section <NUM> is upstream of the last stage variable compressor vane (VCV) <NUM>.

<FIG> is a diagrammatic schematic illustrating the change in flow angle as heat <NUM> is added to the flow in the combustor section <NUM>. The upstream velocity <NUM> has a component in both the circumferential direction and the axial direction. The downstream velocity <NUM>, after heat <NUM> has been added also has components in both the circumferential and axial directions. The component in the circumferential direction remains constant before and after the addition of heat <NUM>. However, the component of the velocity in the axial direction increases as a result of heat <NUM> addition, thereby resulting in a decreased downstream flow angle <NUM>. This is an open swirl embodiment.

<FIG> is a diagrammatic schematic illustrating a confined swirl configuration of the present embodiments, including a plurality of angled vanes <NUM> that guide the airflow through the combustor section <NUM> while heat <NUM> is being added. In one embodiment, the plurality of angled vanes <NUM> may be variable, and may include a plurality of vane actuators <NUM> located upstream of the location where heat <NUM> is added so that the vane actuators <NUM> do not need to withstand such high temperatures. Each vane actuator <NUM> may include a cold pivot <NUM> at the axially upstream end. The plurality of variable vane actuators <NUM> may produce the same effect as variable area turbine nozzles. In other embodiments, the plurality of angled vanes <NUM> may be fixed. In both embodiments, the confined heat addition increases the flow angular momentum such that the desired flow angle as required at the first stage turbine rotor <NUM> is achieved at the outlet of the combustor exit <NUM>. This is a confined swirl embodiment.

<FIG> is a forward-looking-aft view of the combustor section <NUM> according to the present embodiments.

<FIG> is an isometric side view of the combustor section <NUM> according to the present embodiments.

<FIG> is an isometric side view of a gas turbine flow path <NUM> including the combustor inlet inner and outer flow paths, 52A and 52B, the combustor inner and outer flow paths, 54A and 54B, and the combustor exit inner and outer walls, 66A and 66B, respectively. A plurality of angled vanes <NUM> (which may be fixed or variable) circumferentially spaced around the combustor section <NUM> of the gas turbine engine <NUM>. Each angled vane of the plurality of angled vanes <NUM> may be oriented such that they include a component in both the circumferential direction and the axial direction. <FIG> is section view of the gas turbine flow path <NUM>, which is an annular flow path. Stated otherwise, the gas turbine flow path <NUM> in the section view of <FIG> is circumferentially rotated about an engine centerline <NUM>.

<FIG> is an isometric side view of the gas turbine flow path <NUM> of <FIG> including the plurality of angled vanes <NUM>.

<FIG> is an isometric side view of the combustor section <NUM> including the plurality of angled vanes <NUM> spaced circumferentially around the engine centerline <NUM>.

<FIG> is an isometric side view of the combustor section <NUM> including a plurality of angled micro mixer tubes <NUM>. Each of the plurality of angled micro mixers <NUM> disperses fuel and/or a fuel-air mixture into the micro mixer tubes such that heat (not shown) is added to the flow. Each micro mixer <NUM> is a substantially cylindrical tube oriented so as to have a component in the axial direction as well as the circumferential direction. The plurality of angled micro mixers <NUM> are supported by a mounting structure <NUM> extending circumferentially around the combustor cavity <NUM>. Because the micro mixers <NUM> are angled, they produce the desired flow angle at the first stage turbine rotor <NUM> (downstream). Additionally, the orientation of the micro mixers <NUM> may be variable so as to produce the desired flow angle and angular momentum at varying gas turbine operating conditions.

<FIG> is an isometric side view of the combustor section <NUM> including a plurality of angled vanes <NUM>. In a first embodiment, the combustor section <NUM> may include a first fuel injector <NUM> either disposed within the vane <NUM>, or protruding from a radially inner hub <NUM> or via a radially outer casing <NUM>. Furthermore, additional flame holder devices <NUM> may be located between the vanes or as protrusions <NUM> extending from the vanes, or as angled bulkheads <NUM> and air swirlers <NUM> (shown in <FIG>). The angled bulkhead <NUM> may be disposed between adjacent angled vanes <NUM> and may include the swirler <NUM> (shown in <FIG>) and/or a second fuel injector <NUM>.

<FIG> is a side view of the combustor section <NUM> in a corrugated combustor dome arrangement, including a plurality of angled bulkheads <NUM>, each including an air swirler <NUM> circumferentially disposed around a substantially cylindrical fuel injector <NUM>.

<FIG> is an isometric side view of the combustor section <NUM> including a V-gutter style flame holder <NUM> extending circumferentially around the annular combustor cavity <NUM>. As air flows past an outer wing <NUM> and an inner wing <NUM> of the V-gutter style flame holder <NUM>, a vortex forms downstream of the V-gutter style flame holder <NUM> encouraging mixing with fuel dispensed from a center location <NUM> of the V-gutter style flame holder <NUM>.

<FIG> is an isometric side view of the combustor section <NUM> including a V-gutter style flame holder <NUM> and a plurality of angled vanes <NUM> spaced circumferentially around the annular combustor cavity <NUM>.

Each of the embodiments shown in <FIG>, <FIG>, <FIG>, <FIG>, and <FIG> includes a potential mechanism for adjusting the swirl in the combustor section <NUM> to adapt with the operating condition such that the desired downstream flow angle is achieved at the first stage turbine rotor <NUM>.

The confined swirl embodiments of <FIG>, <FIG>, <FIG>, and <FIG> may serve to reduce the required combustor mean radius <NUM>.

The gas turbine engine <NUM>, combustor assembly <NUM> and embodiments described herein reduce the engine length, reduce the engine part count by eliminating the first stage turbine nozzle <NUM> and/or the compressor exit guide vane <NUM>, reduce the cooling flow, reduce losses, and simplify the assembly processes.

Although specific features of various embodiments of the present disclosure may be shown in some drawings and not in others, this is for convenience only.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a compressor section, the compressor section comprising a compressor mean radius; at least one compressor rotor blade and a variable compressor vane (<NUM>);
a combustor section (<NUM>) fluidly coupled downstream of the compressor section, the combustor section comprising a combustor mean radius;
a turbine section fluidly coupled downstream of the combustor section (<NUM>), the turbine section comprising a turbine mid-span radius; and
wherein the combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius,
wherein the variable compressor vane (<NUM>) is disposed downstream of the aft-most compressor rotor blade.