阅读说明：本技术 回流燃烧内衬、回流燃烧器和在其中混合冷却空气的方法 (Recirculation combustion liner, recirculation combustor and method of mixing cooling air therein ) 是由 M·G·吉亚姆布拉 J·D·蒙蒂 S·J·豪厄尔 于 2018-07-25 设计创作，主要内容包括：回流燃烧内衬、回流燃烧器和在回流燃烧器中混合冷却空气的方法,该回流燃烧器包括直线部分、稀释部分和转弯部分。该回流燃烧器接收燃料流,该燃料流点燃并且与冷却空气混合,以形成燃烧气体流。该燃烧气体流通过回流燃烧器行进到发动机的涡轮区段。(A reverse flow combustion liner, a reverse flow combustor, and a method of mixing cooling air in a reverse flow combustor including a straight portion, a dilution portion, and a turn portion. The reverse flow combustor receives a fuel stream that is ignited and mixed with cooling air to form a combustion gas stream. The combustion gas stream travels through a reverse flow combustor to a turbine section of the engine.)

1. A reverse flow combustor for a gas turbine engine, the combustor comprising a combustion liner defining a combustion chamber extending from a dome assembly at a first end and defining a dome height to a turnaround defining a tangential curve, wherein an axial length from the first end to the tangential curve is less than or equal to 150% of the dome height.

2. The reflow burner of claim 1, wherein a set of dilution holes are positioned in a dilution section positioned at an axial distance relative to the first end that is less than or equal to 110% of the dome height to define a second end.

3. The backflow combustor according to claim 2, wherein the tangential curve defines a third end and the dilution portion extends from the second end to the third end.

4. The reflow burner of claim 3, wherein the bend defines a turn portion that extends from the third end to the fourth end and has a cross-sectional area that decreases from the third end to the fourth end.

5. The reflow burner of claim 4, wherein the reduced cross-sectional area continuously decreases at a constant rate.

Technical Field

The present disclosure relates to a reverse flow combustor.

Background

The turbine engine is driven by a flow of air and combustion gases through the engine onto a plurality of rotating turbine blades.

A reverse flow combustor can be provided within the engine for reversing the flow twice, the first for combustion gases (combust the gas) and the second for exhaust gases through the turbine section. Shortening the length of a typical reverse flow combustor is beneficial in terms of cost and engine efficiency on a weight basis.

Disclosure of Invention

In one aspect, the present disclosure is directed to a reverse flow combustion liner comprising: a linear portion having a first end and a second end, the first end defining a dome entrance having a predetermined height; a turn section having a third end and a fourth end, the fourth end defining a turbine inlet. The reverse flow combustion liner further comprises: a dilution portion coupling the second end to the third end and defining a tangential curve at which the dilution portion intersects the turn portion; a set of dilution holes positioned in a dilution portion upstream of a tangential curve, wherein the tangential curve is a distance relative to the combustion dome that is less than or equal to 150% of the dome height.

The dilution portion can have an axial length less than or equal to 20% of the predetermined height. The dilution portion can have an axial length less than or equal to 10% of the predetermined height. The tangential curve can be a distance relative to the dome entrance that is less than or equal to 130% of the predetermined height. The turn section can have a cross-sectional area that decreases from the third end to the fourth end. The reduced cross-sectional area can be continuously reduced. The rate at which the continuous reduction occurs can be constant. A set of dilution holes can be disposed in the combustion liner axially forward relative to the tangential curve. The set of dilution holes can be a plurality of axially spaced sets of dilution holes arranged circumferentially around the liner. The combustion liner can include an inner liner and an outer liner defining a cooling region therebetween. The set of dilution holes can be disposed in the inner liner and the plurality of cooling holes can be disposed on the outer liner.

In another aspect, the present disclosure is directed to a reverse flow combustor for a gas turbine engine, the combustor including a combustion liner defining a combustion chamber extending from a dome assembly at a first end and defining a dome height to a turnaround defining a tangential curve, wherein an axial length from the first end to the tangential curve is less than or equal to 150% of the dome height.

The set of dilution holes can be positioned in a dilution portion positioned at an axial distance less than or equal to 110% of the dome height relative to the first end to define the second end. The tangential curve can define a third end and the dilution portion extends from the second end to the third end. The turn can define a turn portion extending from the third end to the fourth end and the turn portion can have a cross-sectional area that decreases from the third end to the fourth end. The reduced cross-sectional area can continuously decrease at a constant rate.

In yet another aspect, the present disclosure is directed to a method of mixing dilution air in a reverse flow combustor, the method comprising: injecting fuel into the combustion chamber from a dome assembly, the dome assembly having a predetermined dome height; igniting the fuel by an igniter within the combustion chamber to produce combustion gases; and injecting dilution air into the combustion chamber through a set of dilution holes forward of a tangential curve between the curved portion and the straight portion of the combustion chamber, and wherein the tangential curve is a distance relative to the combustor dome that is less than or equal to 150% of the dome height.

The method can also include mixing the cooling air with the combustion gases in the turning portion of the combustion chamber. A set of dilution holes can be positioned within the dilution portion, wherein the dilution portion has an axial length less than or equal to 20% of the predetermined height relative to the tangential curve.

Technical solution 1. a reverse flow combustion liner, comprising:

a linear portion having a first end and a second end, the first end defining a dome entrance having a predetermined height;

a turn section having a third end and a fourth end, the fourth end defining a turbine inlet;

a dilution portion coupling the second end to the third end and defining a tangential curve at which the dilution portion intersects the turn portion;

a set of dilution holes positioned in a dilution portion upstream of the tangential curve, wherein the tangential curve is a distance relative to the dome inlet that is less than or equal to 150% of the predetermined height.

The reverse flow combustion liner of claim 1, wherein the dilution portion has an axial length less than or equal to 20% of the predetermined height.

Claim 3. the reverse flow combustion liner of claim 1, wherein the dilution portion has an axial length less than or equal to 10% of the predetermined height.

Claim 4. the reverse flow combustion liner of claim 1, wherein the tangential curve is a distance relative to the dome inlet that is less than or equal to 130% of the predetermined height.

Claim 5. the reverse flow combustion liner of claim 1, wherein the turn portion has a cross-sectional area that decreases from the third end to the fourth end.

Claim 6. the reverse flow combustion liner of claim 5, wherein the reduced cross-sectional area continuously decreases.

Claim 7. the reverse flow combustion liner of claim 6, wherein the rate at which the continuous decrease occurs is constant.

The reverse flow combustion liner of claim 1, wherein the set of dilution holes are disposed in the combustion liner axially forward relative to the tangential curve.

The reverse flow combustion liner of claim 1, wherein the set of dilution holes is a plurality of axially spaced sets of dilution holes arranged circumferentially around the liner.

Claim 10. the reflowed combustion liner of claim 1, wherein the combustion liner comprises an inner liner and an outer liner defining a cooling region therebetween.

The reverse flow combustion liner of claim 10, wherein the set of dilution holes is disposed in the inner liner and a plurality of cooling holes is disposed on the outer liner.

Solution 12. a flashback combustor for a gas turbine engine, said combustor comprising a combustion liner defining a combustion chamber extending from a dome assembly at a first end and defining a dome height to a turn defining a tangential curve, wherein an axial length from said first end to said tangential curve is less than or equal to 150% of said dome height.

The reverse flow combustor of claim 12, wherein a set of dilution holes are positioned in a dilution section positioned at an axial distance relative to the first end that is less than or equal to 110% of the dome height to define a second end.

Claim 14 the recirculation burner of claim 13, wherein the tangential curve defines a third end and the dilution portion extends from the second end to the third end.

Claim 15 the reverse flow combustor of claim 14, wherein the turn defines a turn portion that extends from the third end to the fourth end and the turn portion has a cross-sectional area that decreases from the third end to the fourth end.

Claim 16. the reverse flow combustor of claim 15, wherein the reduced cross-sectional area continuously decreases at a constant rate.

A method of mixing cooling air in a reverse flow combustor, the method comprising:

injecting fuel into a combustion chamber from a dome assembly, the dome assembly having a predetermined dome height;

igniting fuel in the combustion chamber by an igniter to produce combustion gases; and

injecting cooling air into the combustion chamber through a set of dilution holes forward of a tangential curve between a turnaround portion and a straight portion of the combustion chamber, and wherein the tangential curve is a distance relative to the dome assembly that is less than or equal to 150% of the predetermined dome height.

Claim 18 the method of claim 17, further comprising mixing cooling air with combustion gases in the turning section of the combustion chamber.

Solution 19. the method of solution 18, wherein the set of dilution holes is positioned within the dilution portion, wherein the dilution portion has an axial length relative to the tangential curve that is less than or equal to 20% of the predetermined height.

Drawings

In the drawings:

FIG. 1 is a schematic cross-sectional view of a free-turbine engine for an aircraft having a reverse flow combustor.

Fig. 2 is an enlarged cross-sectional view of a reverse flow combustor from the prior art.

FIG. 3 is an enlarged cross-sectional view of a reverse flow combustor for the free turbine engine of FIG. 1.

FIG. 4 is another cross-sectional view of the recirculation burner of FIG. 3, wherein different burner widths are shown.

Detailed Description

Aspects of the present disclosure described herein relate to a reverse flow combustor. For purposes of illustration, the present disclosure will be described with respect to a free turbine engine. However, it should be understood that the aspects of the present disclosure described herein are not so limited and the reverse flow combustor as described herein can be implemented in other turbine engines, including but not limited to turbojet engines, turboprop engines, turboshaft engines, and turbofan engines. Aspects of the disclosure discussed herein may have general applicability in non-aircraft engines having a reverse flow combustor, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

As used herein, the term "forward" or "upstream" refers to moving in a direction toward the engine inlet, or one component being relatively closer to the engine inlet than the other component. The term "aft" or "downstream" as used in conjunction with "forward" or "upstream" refers to a direction toward the outlet of the engine or relatively closer to the engine outlet than another component. Further, as used herein, the term "radial" or "radially" refers to a dimension extending between a central longitudinal axis of the engine and an outer engine circumference. Furthermore, as used herein, the term "group" or a "set" of elements can be any number of elements, including only one.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, forward, rearward, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, rearward, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of aspects of the disclosure in relation to this specification. Unless specified otherwise, connection references (e.g., attached, coupled, connected, and engaged) are to be construed broadly and may include intermediate members between a series of elements and relative movement between elements. Thus, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for illustrative purposes only, and the dimensions, positions, order, and relative sizes reflected in the drawings of the present disclosure may vary.

Referring to FIG. 1, an engine 10 has a generally longitudinally extending axis or centerline 12 extending from a forward portion 14 to an aft portion 16. Engine 10 is a free turbine engine wherein air is drawn in at inlet 18 and flows generally aft 16 in the direction of forward 14. Engine 10 includes in downstream series flow relationship: a compressor section 22 including an inlet 18; a compressor 24 and a centrifugal compressor 26; a combustion section 28 including a combustor 30; a turbine section 32 including a HP turbine 34 and a LP turbine 36; an exhaust section 38; and a propeller section 40. The propeller section 40 includes a propeller hub 42 having a plurality of propeller blades 44 radially arranged about the centerline 12. The engine housing 46 can form an annular housing for each section.

An HP shaft (draft) or spool 48, coaxially disposed about the centerline 12 of the engine 10, drivingly connects the HP turbine 34 to the compressor 24 and the centrifugal compressor 26. An LP shaft or spool 50, which is collinear with and separate from the HP spool 48, and coaxially disposed about the centerline 12 of the engine 10, drivingly connects the LP turbine 36 to the propeller hub 42. The shafts 48, 50 are rotatable about the engine centerline 12 and are coupled to a plurality of rotatable elements that may collectively define a rotor 51.

The compressor 24 includes at least one compressor stage 54 in which a set of compressor blades (blades) 56 rotate relative to a corresponding set of static compressor vanes (vane)60 (also referred to as nozzles) to compress or pressurize a fluid flow through the stage. The centrifugal compressor 26 can include an impeller 52 having a set of impeller blades 58. In a single compressor stage 54, a plurality of compressor blades 56 may be provided in a ring and may extend radially outward from the blade platform to the blade tip relative to the centerline 12, while corresponding static compressor vanes 60 are positioned upstream of and adjacent to the rotating blades 56. It should be noted that the number of blades, vanes, and compressor stages shown in FIG. 1 is chosen for illustrative purposes only, and other numbers are possible.

The vanes 56, 58 for each stage of the compressors 24, 26 are mounted to a disc 61 that is mounted to the corresponding HP spool 48, with each stage having its own disc 61. The buckets 60 for the stages of the compressor may be mounted to the engine casing 46 in a circumferential arrangement.

The HP and LP turbines 34, 36 each include a plurality of turbine stages 64, 66, with a set of turbine blades 68, 70 rotating relative to a corresponding set of stationary turbine buckets 72, 74 (also referred to as nozzles) to extract energy from the fluid flow through the stages. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 may be provided in a ring and may extend radially outward relative to the centerline 12 from the blade platform to the blade tip, while corresponding static turbine buckets 72, 74 are positioned upstream of and adjacent to the rotating blades 68, 70. It should be noted that the number of blades, buckets, and turbine stages shown in FIG. 1 is chosen for illustrative purposes only, and other numbers are possible.

The blades 68 for the HP turbine stages 64 are mounted to a disc 71 that is mounted to the corresponding HP spool 48, with each stage having a dedicated disc 71. The blades 70 for the LP turbine stages 66 are mounted to a disk 73 that fits into the corresponding LP spool 50, with each stage having a dedicated disk 73. The buckets 72, 74 for the respective stages of the turbines 34, 36 may be mounted to the core casing 46 in a circumferential arrangement.

Complementary to the rotor portions, the stationary portions of the engine 10, such as the static vanes 60, 72, 74 in the compressor section 22 and the turbine section 32, are also referred to individually or collectively as the stator 63. Thus, the stator 63 may refer to a combination of non-rotating elements throughout the engine 10.

In operation, the flow of air entering the inlet 18 is channeled into the compressor 24, which in turn supplies pressurized air 76 to the centrifugal compressor 26, which further pressurizes the air. The pressurized air 76 from the centrifugal compressor 26 is mixed with fuel in the combustor 30, where the fuel is combusted, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, driving the compressor 24 and the centrifugal compressor 26. The HP turbine 34 discharges the combustion gases into the LP turbine 36, which extracts additional work to drive the propeller hub 42, and the exhaust gases are ultimately discharged from the engine 10 via the exhaust section 38.

A portion of the pressurized gas stream 76 may be drawn from the compressor section 22 as bleed air 77. Bleed air 77 may be drawn from pressurized airflow 76 and used in other areas of the engine or aircraft, such as being provided to engine components that require cooling. The temperature of the pressurized gas stream 76 entering the combustor 30 is significantly elevated relative to the temperature of the gas stream entering at the inlet 18. Thus, the cooling provided by bleed air 77 is necessary for operation of these engine components in a high temperature environment.

The LP turbine 36 drives the LP spool 50 to rotate the propeller hub 42 and subsequently the propeller blades 44. Reduction gearbox 80 is disposed between LP spool 50 and propeller hub 42. In addition, an accessory gearbox 82 can be provided to power other components of the engine 10, including by way of non-limiting example, a fuel pump, a fuel controller, an oil pump, a starter/generator, and a tachometer.

Fig. 2 shows a cross-sectional view of a reflow burner 30 according to the prior art. The reverse flow combustor 30 is disposed radially outward relative to the HP turbine 34 and is disposed axially upstream and forward of the LP turbine 36. The reverse flow combustor 30 may include a dome assembly 102 and an annular liner assembly 100 that define a combustion chamber 103, and a plurality of fuel injectors 104 fluidly coupled to the combustion chamber 103. The reverse flow combustor 30 is disposed within the combustor casing 106.

An interior 108 of the annular liner assembly 100 may surround the HP turbine 34. The outer portion 110 of the annular liner assembly 100, together with the combustor casing 106, defines at least a portion of a compressed air passage 112 through which compressed air 113 can flow from the compressor section 22 to the reverse flow combustor 30. A plurality of dilution holes 114 may be provided in annular liner assembly 100, allowing at least a portion of compressed air 113 to pass from compressed air passage 112 to combustion chamber 103.

A plurality of fuel injectors 104 are coupled to and disposed within the dome assembly 102 at a dome entrance 105 that includes a flared cone 107. The dome entrance 105 defines a Dome Height (DH). The plurality of fuel injectors 104 include a fuel inlet 116 adapted to receive a flow of fuel 118, a fuel outlet 120 in fluid communication with the combustion chamber 103, and a linear fuel passage 122 extending therebetween. It should be appreciated that one or more fuel injectors 104 can have different characteristics, and the plurality of fuel injectors 104 is for illustrative purposes only and is not intended to be limiting. A swirler 124 can be disposed at the dome inlet 105 to approximate the incoming air to the fuel swirl exiting the fuel injector 104 and provide a uniform mixture of air and fuel entering the combustor.

Compressed air 113 and fuel 118 are supplied to combustion chamber 103 through dome inlet 105 and are subsequently ignited within combustion chamber 103 by one or more igniters 126, thereby generating combustion gases 128. Combustion gases 128 are mixed using additional air supplied through a plurality of dilution holes 114 disposed in inner portion 108 and outer portion 110 of annular liner assembly 100 of combustion chamber 103 and mixed in a dilution portion 132 of combustion chamber 103, after which combustion gases 128 flow through turn portion 134 and are discharged into HP turbine 34. The first radius R1 and the second radius R2 intersect the inner portion 108 and the outer portion 110 at the tangential curve 162 such that the curve is defined as a loop line at which the dilution portion 132 intersects the turn portion 134. The dilution portion 132 can also be referred to as a dilution zone or a mixing zone and has an axial length (DL). The gases discharged from HP turbine 34 are then channeled into LP turbine 36. The tangential curve 162 defines the start of the dilution portion 132. The Axial Length (AL) measured from the flared cone 107 to the tangential curve 162 is at least 180% of the Dome Height (DH).

Fig. 3 illustrates a reverse flow combustor 130 according to a preferred aspect of the present disclosure described herein. The recirculation burner 130 is similar to the recirculation burner 30 of the prior art and therefore like components will be designated by reference numerals increased by 100. It should be understood that the description of similar components of the reverse flow combustor 30 applies to the reverse flow combustor 130 unless otherwise noted.

The annular liner assembly 200 and the dome assembly 202 define a combustion chamber 203, and a plurality of fuel injectors 204 are fluidly coupled to the combustion chamber 203. The annular liner assembly 200 can include an inner liner 250 and an outer liner 252 defining a cooling region 254 therebetween. A plurality of cooling holes 256 can be disposed throughout outer liner 252.

A set of dilution holes 214 is disposed in the inner liner 250 and may also be disposed in the outer liner 252 such that the set of dilution holes 214 surrounds the compressed air passage 212 and fluidly couples the compressed air passage 212 with the combustion chamber 203. The set of dilution holes 214 can be one or more rows of dilution holes 214 axially spaced from one another and arranged annularly about the annular liner assembly 200. In a preferred aspect of the present disclosure, the set of dilution holes 214 is a single row of dilution holes, however, other smaller or "trim" holes may be applied in other rows if desired. It should be understood that the annular liner assembly 200 can be a single liner, wherein the single annular liner includes a set of dilution holes 214.

The annular liner assembly 200 includes a straight portion 260, a dilute portion 232, and a turn portion 234. The linear portion 260 extends from a first end (a) at the dome assembly 202 defined by the flared cone 207. The straight portion 260 can also be referred to as a main area or firing area. The dome assembly 202 defines a dome entrance 205 having a predetermined height or dome height (H) defining a diameter at a first end (a). The linear portion 260 terminates at a second end (B) at a point directly forward of the plane of the dilution holes 214 in the dilution portion 232 to define an axial length (S) relative to the flared cone 207 equal to 100% of the dome height (H). It is contemplated that the axial length (S) of the linear portion 260 can be between 90% and 130% of the dome height (H).

The second end (B) is positioned directly forward of the set of dilution holes 214 such that the dilution holes 214 are located within the dilution portion 232. The dilution portion 232 defines an axial length (T) that is equal to or less than 30% of the dome height (H). It is also contemplated that the axial length (T) of the dilution portion 232 can be 20% of the dome height (H). The dilution portion 232 extends from the second end (B) to the third end (C) to define an axial length (T).

The third end (C) defines a tangential curve 262 from which the turn portion 234 begins. The first radius R1 and the second radius R2 intersect the inner portion 208 and the outer portion 210 at the tangential curve 262 such that the curve is defined as a loop line at which the dilution portion 232 intersects the turn portion 234. The tangential curve 262 can be positioned at an axial distance (S + T) of 130% of the dome height (H) as measured from the flared cone 207. It is contemplated that the axial length (S + T) can be between 110% and 150% of the dome height (H).

The turn portion 234 extends from a third end (C) to a fourth end (D) defining a burner outlet 264. The combustor outlet 264 transitions to a turbine inlet 266 (FIG. 1) of the HP turbine 34. Turning portion 234 redirects fuel 218 from a primarily aft direction within combustion chamber 203 to a primarily forward direction at HP turbine 34.

It should be understood that the axial measurements made are from exemplary locations and are not limited to measurements relative to the flare cone 207 and the tangential curve 262. It is contemplated that the axial measurement can be taken from any location within the dome assembly 202 and any point within the combustion chamber 203 that defines the beginning of the turn. The axial length (T) of the dilution portion 232 can be any length from 10% to 30% of the dome height (H) as described herein.

Referring to fig. 4, a cross-sectional view of the flashback chamber 130 shows a turn portion 234 having a continuously decreasing cross-sectional area (CA) extending from a third end (C) to a fourth end (D). The rate at which the cross-sectional area (CA) decreases may be a constant rate; by way of non-limiting example, CA2 (i.e., the second cross-sectional area) is 7% less than CA1 (i.e., the first cross-sectional area) and CA3 (i.e., the third cross-sectional area) is also 7% less than CA2, and so on. The velocity need not be constant and the difference between successive cross-sectional areas lies within 10 degrees of each other with respect to the radial plane and may be 10% or less smaller than the preceding cross-sectional area (CA).

The method of mixing dilution air 268 in the reverse flow combustor 130 as described herein includes injecting fuel 218 from the dome assembly 202 into the combustion chamber 203. The fuel 218 is ignited within the combustion chamber 203 by an igniter 226 to produce combustion gases 228. The method then includes injecting dilution air 268 into the combustion chamber 203 through a set of dilution holes 214 in the dilution section 232 forward of a tangential curve 262 between the turning section 234 and the straight section 260 of the combustion chamber 203. The axial length (T) of the dilution portion 232 is much smaller than the axial length (DL) of the dilution portion 132 of the combustion chamber 103 (FIG. 2). The method thus further includes mixing dilution air 268 with the combustion gases 228 in the turn portion 234 of the combustion chamber 203.

Aspects of the present disclosure described herein relate to a reverse flow combustor designed to minimize the overall length of combustion gas flow within the combustion chamber. The combustor includes a conventional straight section or main area, and a short dilution section or dilution area. Providing a single row of openings directly forward of the tangential curve to the turning section takes advantage of the inherent trajectory that dilution air follows when introduced and maximizes the amount of mixing of combustion gases with dilution air in the small dilution section. Other mixing still occurs in the turn section and is complete before being introduced into the turbine section.

Conventional reverse flow combustors have turn sections that require additional cooling surface area that requires air that would otherwise be used for dilution and combustion. Such increased cooling in conventional reverse flow combustors requires that the combustor be limited to a lower pressure ratio of 12:1 or less. Typically, dilution and mixing is done before the gases enter the turn duct, so that the length of the burner before the turn section is similar to that of a comparable axial burner. The reverse flow combustor as described herein allows the final stage of dilution mixing to occur in the turn section, thus allowing for an overall shorter combustion chamber and thus reduced cooling surface area requirements.

Benefits associated with a reverse flow combustor as described herein include reducing the cooling surface area of the combustor to minimize the cooling air required to keep the liner temperature within acceptable limits, while still allowing sufficient main area volume for flame holding and ignition. There are also engine weight and packaging benefits realized by providing an overall shorter combustion liner and casing. The savings in cooling air requirements allow the use of the recuperator for higher pressure ratio cycles (by way of non-limiting example, 16: 1 pressure ratio cycles).

When described with reference to a free turbine propeller engine, it should be appreciated that the reverse flow combustor as described herein can be any engine having a high pressure ratio engine cycle. It should be appreciated that the application of the various aspects of the present disclosure discussed herein is not limited to turbine engines having a propeller section, but is also applicable to engines having a fan and booster section as well as turbojet and turbine engines.

To the extent not already described, the different features and structures of the various embodiments may be combined with or substituted for one another as desired. A feature that is not illustrated in all embodiments is not meant to be construed as such, but is done so for simplicity of description. Thus, the various features of the different embodiments can be mixed and matched as necessary to form new embodiments, whether or not such new embodiments are explicitly described. The present disclosure encompasses all combinations or permutations of features described herein.

This written description uses examples to describe aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the aspects of the disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.