Supplemental thrust system with rotating detonation combustor

An assembly is provided for a turbine engine. This turbine engine assembly includes a supplemental thrust section and a duct. The supplemental thrust section includes a rotating detonation combustor. The duct includes a supplemental thrust section inlet fluidly coupled with and leading to the rotating detonation combustor. The supplemental thrust section inlet has a flow area that decreases as at least a first portion of the supplemental thrust section inlet extends towards the rotating detonation combustor.

BACKGROUND OF THE DISCLOSURE

1. Technical Field

This disclosure relates generally to a turbine engine and, more particularly, to a supplemental thrust system for the turbine engine.

2. Background Information

Various types and configurations of gas turbine engines are known in the art for propelling aircraft. These known turbine engines may be configured with relatively large aircraft as well as with relatively small aircraft. Typically, turbine engines configured with small aircraft have relatively low thrust ratings. Such small aircraft therefore may have difficultly reaching and/or maintaining high Mach flight; e.g., supersonic flight. While it is known to provide some large turbine engines with a supplemental thrust system, these known supplemental thrust systems are relatively bulky and may be difficult to implement with small/compact gas turbine engines. There is a need in the art therefore for a supplemental thrust system for gas turbine engines, particularly relatively small/compact gas turbine engines.

SUMMARY OF THE DISCLOSURE

According to an aspect of the present disclosure, an assembly is provided for a turbine engine. This turbine engine assembly includes a supplemental thrust section and a duct. The supplemental thrust section includes a rotating detonation combustor. The duct includes a supplemental thrust section inlet fluidly coupled with and leading to the rotating detonation combustor. The supplemental thrust section inlet has a flow area that decreases as at least a first portion of the supplemental thrust section inlet extends towards the rotating detonation combustor.

According to another aspect of the present disclosure, another assembly is provided for a turbine engine. This turbine engine assembly includes a supplemental thrust section, a duct and a bypass passage. The supplemental thrust section includes a rotating detonation combustor. The duct is configured to direct gas into the supplemental thrust section during a first mode. The bypass passage is configured, during a second mode, to: receive at least a portion of the gas form the duct; and direct the portion of the gas to bypass the supplemental thrust section.

According to still another aspect of the present disclosure, another assembly is provided for a turbine engine. This turbine engine assembly includes a compressor section, a combustor section, a turbine section and a supplemental thrust section arranged sequentially along a flowpath between an engine inlet and an engine exhaust. The supplemental thrust section includes a rotating detonation combustor. The turbine engine assembly also includes a bypass passage. The bypass passage is configured to: receive gas from the flowpath upstream of the combustor section; and provide the gas to the supplemental thrust section during at least a mode of operation of the rotating detonation combustor.

The bypass passage may direct the gas into a combustion chamber of the rotating detonation combustor.

The supplemental thrust section may be operational during the first mode. The supplemental thrust section may be non-operational during the second mode.

The supplemental thrust section may be operational during the first mode and the second mode.

The turbine engine assembly may also include an exhaust nozzle fluidly coupled with and downstream of the supplemental thrust section and the bypass passage.

The turbine engine assembly may also include a turbine section upstream of the supplemental thrust section. The supplemental thrust section inlet may fluidly couple the turbine section to the rotating detonation combustor.

The supplemental thrust section inlet may be configured as or otherwise include an annular supplemental thrust section inlet.

The supplemental thrust section inlet may include an inner wall and an outer wall. The inner wall may extend circumferentially about a centerline. The inner wall may extend radially outward towards the outer wall as the first portion of the supplemental thrust section inlet extends axially along the centerline towards the rotating detonation combustor.

The supplemental thrust section inlet may include an inner wall and an outer wall. The outer wall may extend circumferentially about a centerline. The outer wall may extend radially inward towards the inner wall as the first portion of the supplemental thrust section inlet extends axially along the centerline towards the rotating detonation combustor.

The flow area may increase as a second portion of the supplemental thrust section inlet extends away from the first portion of the supplemental thrust section inlet and towards the rotating detonation combustor.

The supplemental thrust section inlet may form a convergent-divergent diffuser inlet to the supplemental thrust section.

The turbine engine assembly may also include a variable area exhaust nozzle fluidly coupled with and configured downstream of the supplemental thrust section.

The turbine engine assembly may also include a turbine section, a combustor section and a bypass passage. The turbine section may be fluidly coupled with and upstream of the duct. The combustor section may be fluidly coupled with and upstream of the turbine section. The bypass passage may be configured to provide gas to the supplemental thrust section that bypasses the combustor section and the turbine section.

The turbine engine assembly may also include a compressor section fluidly coupled with and upstream of the combustor section. The bypass passage may be configured to receive the gas from the compressor section.

The turbine engine assembly may also include a turbine section and a flow regulator. The turbine section may be fluidly coupled with and upstream of the duct. The flow regulator may be configured to regulate a flow of gas exhausted from the turbine section into the supplemental thrust section.

The flow regulator may be configured as or otherwise include a variable area diffuser.

The flow regulator may be configured as or otherwise include a moveable sleeve.

The flow regulator may be configured as or otherwise include an array of variable stator vanes.

The flow regulator may be configured as or otherwise include a translating plug.

The turbine engine assembly may also include a bypass passage configured to direct at least a portion of gas flowing within the duct to bypass the supplemental thrust section. The flow regulator may be configured as or otherwise include a valve configured with the bypass passage.

The turbine engine assembly may also include a bypass passage configured to direct at least a portion of gas flowing within the duct around the supplemental thrust section.

The bypass passage may be configured as or otherwise include an annular inner duct arranged inward of and extending longitudinally along the rotating detonation combustor. The bypass passage may also or alternatively be configured as or otherwise include an annular outer duct arranged outward of and extending longitudinally along the rotating detonation combustor.

DETAILED DESCRIPTION

FIG.1is a schematic illustration of a gas turbine engine20; e.g., a turbojet engine. This turbine engine20extends along an axial centerline22(e.g., a rotational axis) between an upstream airflow inlet24(e.g., an engine inlet) and a downstream exhaust outlet26(e.g., an engine exhaust). The turbine engine includes a compressor section28, a combustor section29, a turbine section30, a supplemental thrust section31with a supplemental thrust system, and an exhaust section32.

The engine sections28-32are arranged sequentially along the axial centerline22within an engine housing34. This engine housing34includes an engine case36that houses one or more of the engine sections28-32. The engine case36ofFIG.1, for example, extends axially along the axial centerline22from the airflow inlet24to an exhaust nozzle38at the exhaust outlet26.

Each of the engine sections28and30includes a respective rotor40and42. Each of these rotors40and42includes a plurality of rotor blades arranged circumferentially around and connected to one or more respective rotor disks. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s).

The compressor rotor40is connected to the turbine rotor42through an engine shaft44. This engine shaft44is rotatably supported by a plurality of bearings (not shown), which bearings are connected to the engine housing34by at least one stationary structure such as, for example, an annular support strut (not shown).

During operation, air enters the turbine engine20through the airflow inlet24. This air is directed into at least a core flowpath46. This core flowpath46extends sequentially through the engine sections28-32to the exhaust outlet26. The air within the core flowpath46may be referred to as core air.

The core air is compressed by the compressor rotor40and directed into a combustion chamber48of a combustor50in the combustor section29. Fuel is injected into the combustion chamber48and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof flow through and cause the turbine rotor42to rotate. The rotation of the turbine rotor42drives rotation of the compressor rotor40and, thus, compression of the air received from the airflow inlet24. The combustion products are directed out of the turbine engine20through the exhaust nozzle38to provide forward engine thrust.

The turbine engine20may be configured to operate in various different modes of operation such as, but not limited to, low thrust operation and high thrust operation. The turbine engine20may be operated in low thrust operation where, for example, additional/supplemental forward engine thrust provided by the supplemental thrust section31is not required. Examples of low thrust operation include, but are not limited to, aircraft takeoff, low speed aircraft cruise, subsonic aircraft flight and/or aircraft landing. On the other hand, the turbine engine20may be operated in high thrust operation where, for example, additional/supplemental forward engine thrust provided by the supplemental thrust section31is required or otherwise beneficial to aircraft flight. Examples of high thrust operation include, but are not limited to, high speed aircraft cruise, supersonic aircraft flight and/or high speed aircraft maneuvers.

When the supplemental thrust section31is non-operation, gas (e.g., the combustion products) exhausted from the turbine section30may flow through the supplemental thrust section31substantially uninterrupted (e.g., without introduction of additional fuel, without additional combustion, etc.) to the exhaust nozzle38. Thus, the supplemental thrust section31does not provide any forward engine thrust to supplement the forward engine thrust already provided by the exhaust gas. On the other hand, when the supplemental thrust section31is operational, additional fuel is mixed with the exhaust gas within the supplemental thrust section31. This fuel-gas mixture is ignited and combustion products thereof flow out of the supplemental thrust section31, through the exhaust section32, to the exhaust nozzle38. Thrust provided by the additional combustion products supplement the thrust already provided by the combustion products exhausted from the turbine section30and, thus, provide the turbine engine20with additional forward engine thrust. The level of the additional forward engine thrust is related to a quantity of additional fuel mixed with the exhaust gas and ignited within the supplemental thrust section31. The fuel may be injected into the supplemental thrust section31at a steady, uniform flow rate, or may be varied to control the level of the additional forward engine thrust.

The supplemental thrust section31ofFIG.1includes a rotating detonation combustor52, which may also be referred to as a rotating detonation engine. This rotating detonation combustor52may have a relatively radially and/or axially compact form.

The rotating detonation combustor52is configured to inject the additional fuel into the core flowpath46for mixing with the gas exhausted from the turbine section30. This fuel-gas mixture is ignited within the rotating detonation combustor52to generate detonation waves, where each of the detonation waves travels circumferentially around and axially through a rotating detonation combustor (RDC) combustion annulus54of the rotating detonation combustor52. By contrast, combustion products within a traditional combustor/supplemental thrust system flow substantially axially therethrough.

Referring toFIG.2, the rotating detonation combustor52is configured with one or more internal volumes. The rotating detonation combustor52ofFIG.2, for example, includes a rotating detonation combustor (RDC) combustion chamber56and the combustion annulus54. The rotating detonation combustor52also includes a fuel injection system58and an igniter system60.

The combustion chamber56ofFIGS.2and3is configured as a tubular chamber. Referring toFIG.2, this combustion chamber56extends radially between and is formed by a radial inner wall62and a radial outer wall64. Each of these chamber walls62and64extends circumferentially about (e.g., completely around) the axial centerline22. Each of the chamber walls62and64extends axially along the axial centerline22from (or about) a supplement thrust section (STS) inlet66to (or about) an upstream end of the combustion annulus54. The combustion chamber56ofFIG.2is configured with a uniform cross-sectional flow area along its longitudinal length; e.g., along the axial centerline22. However, in other embodiments, the cross-sectional flow area of the combustion chamber56may be non-uniform; e.g., converging and/or diverging as the chamber56extends axially downstream.

The combustion annulus54ofFIGS.2and4is configured as a tubular chamber. Referring toFIG.2, this combustion annulus54extends radially between and is formed by a radial inner wall68and a radial outer wall70. Each of these annulus walls68and70extends circumferentially about (e.g., completely around) the axial centerline22. Each of the annulus walls68and70extends axially along the axial centerline22from (or about) a downstream end of the combustion chamber56to (or about) an upstream end of the exhaust section32. The combustion annulus54ofFIG.2is configured with a uniform cross-sectional flow area along its longitudinal length; e.g., along the axial centerline22. However, in other embodiments, the cross-sectional flow area of the combustion annulus54may be non-uniform; e.g., converging and/or diverging as the annulus54extends axially downstream.

The annulus inner wall68may be radially aligned with, connected to and/or formed integral with the chamber inner wall62. However, in other embodiments, the inner walls62and68may be discrete and/or separated from one another by another section of wall, a flow impediment, a radial jog, etc. The annulus outer wall70may be radially aligned with, connected to and/or formed integral with the chamber outer wall64. However, in other embodiments, the outer walls64and70may be discrete and/or separated from one another by another section of wall, a flow impediment, a radial jog, etc.

The fuel injection system58is configured to inject fuel into the combustion chamber56. The fuel injection system58ofFIGS.2and3, for example, includes one or more fuel injectors72. The fuel injectors72ofFIG.3are arranged circumferentially about the axial centerline22in an annular array. Each fuel injector72is configured to spray (e.g., an atomized spray) of the fuel into the combustion chamber56. Each fuel injector72ofFIG.2, for example, is configured with a nozzle orifice74through the chamber outer wall64, where that fuel injector72is operable to inject the fuel radially inward through the respective nozzle orifice74into the combustion chamber56. The present disclosure, however, is not limited to such an exemplary fuel injector configuration. For example, referring toFIG.5, each of the fuel injectors72may alternatively be configured with its nozzle orifice74through the chamber inner wall62, where that fuel injector72is operable to inject the fuel radially outward through the respective nozzle orifice74into the combustion chamber56. In another example, referring toFIG.6, each of the fuel injectors72may alternatively be configured with a vane array76within the core flowpath46. This vane array76may include one or more vanes78arranged circumferentially about the axial centerline22in an annular array (one vane78visible inFIG.6), where each vane78extends radially across the core flowpath46/the combustion chamber56. Each fuel injector72is arranged with a respective one of the vanes78, and is operable to inject the fuel laterally (e.g., circumferentially or tangentially) out of one or more respective nozzle orifices74into the combustion chamber56. Of course, the fuel injection system58may alternatively include any combination of the fuel injectors72shown inFIGS.2,5and/or6, and/or various other types of fuel injectors72capable of delivering the fuel to the combustion chamber56. Furthermore, while the fuel injection system58is described as including fuel injectors72above, one or more of these fuel injectors72may be replaced with an orifice79in the respective component (e.g.,64,62or78as shown, for example, inFIG.7. This orifice79ofFIG.7fluidly coupled with an adjacent fuel passage81of, for example, a fuel manifold83.

Referring toFIG.2, the igniter system60includes one or more igniters80; e.g., fuel detonators. These igniters80may be arranged circumferentially about the axial centerline22in a similar manner as described above with respect to the fuel injectors72. Each of the igniters80is arranged downstream of the fuel injectors72. Each of the igniters80is configured to ignite a mixture of the fuel and the exhaust gas.

During operation of the supplemental thrust section31ofFIG.2, the STS inlet66directs the gas exhausted from the turbine section30(seeFIG.1) into the combustion chamber56. The fuel injection system58inject the fuel into the combustion chamber56for mixing with the exhaust gas. This fuel-gas mixture is ignited by the igniter system60, which initiates detonation of the fuel-gas mixture. Referring toFIGS.8and9, the detonation may correspond to an ignition or combustion of the fuel-gas mixture at a particular location about a circumference of the combustion annulus54.

The detonation may continuously travel around a circumference of the combustion annulus54. For example, a detonation79may initially occur at a starting location81and may then propagate/travel in a circumferential direction82. A first location84within the rotating detonation combustor52, preceding the detonation79, may include a relatively large density of the fuel-gas mixture86. As the detonation79reaches this first location84, the density of the fuel-gas mixture86facilitates detonation thereof. Following this detonation79, the fuel-gas mixture86may be burned away and a force of the detonation79may temporarily resist entry of additional fuel-gas mixture into the rotating detonation combustor52. Thus, a second location88that has recently detonated may have a relatively low density of the fuel-gas mixture86. The detonation79may thereby continue to rotate about the axial centerline22through the rotating detonation combustor52in the circumferential direction82. One skilled in the art, of course, will recognize there may be certain variations in the foregoing detonation dynamics.

To facilitate rotating detonation combustor operation within the supplemental thrust section31, the turbine engine20of the present disclosure is configured to condition the exhaust gas output from the turbine section30. In particular, referring toFIG.1, the gas exhausted by the turbine section30and directed through the core flowpath46via an inter-turbine-supplemental thrust section duct90towards the supplemental thrust section31may be relatively turbulent, high temperature, high speed and oxygen deficient. The exhaust gas received by the supplemental thrust section31may thereby be relatively vitiated, particularly compared to the relatively low temperature, low speed and oxygen rich core air received by the combustor in the combustor section29. To accommodate arrangement of the rotating detonation combustor52downstream of the turbine section30, the exhaust gas provided to the rotating detonation combustor52via the duct90may be conditioned using one or more techniques such as, but not limited to:(1) Accelerating the exhaust gas provided to the rotating detonation combustor52;(2) Decelerating the exhaust gas provided to the rotating detonation combustor52;(3) Diffusing the exhaust gas provided to the rotating detonation combustor52;(4) Providing supplemental oxygen-rich gas to the rotating detonation combustor52; and/or(5) Regulating gas flow to and/or through the rotating detonation combustor52. The rotating detonation combustor52may thereby be subject to flow conditions closer to that experienced by a combustor in a typical inter-compressor-turbine location; e.g., the combustor50ofFIG.1.

Referring toFIG.10, the STS inlet66may be configured to accelerate the gas exhausted from the turbine section30towards/into the combustion chamber56. The STS inlet66ofFIG.10, for example, is configured as a convergent inlet nozzle, which inlet nozzle is axially between and fluidly couples the turbine section30and the combustion chamber56. The STS inlet66may be configured as a part of the duct90and/or the supplemental thrust section31.

The STS inlet66ofFIG.10includes a radial inner wall92and a radial outer wall94. The inlet inner wall92extends circumferentially about (e.g., completely around) the axial centerline22. At least a portion (or an entirety) of the inlet inner wall92extends radially outward towards the inlet outer wall94as the STS inlet66extends axially downstream towards (or to) the combustion chamber56and its inner wall62. The inlet inner wall92ofFIG.10may thereby be canted relative to the axial centerline22, and converge radially outwards towards the inlet outer wall94. Similarly, the inlet outer wall94extends circumferentially about (e.g., completely around) the axial centerline22. The inlet outer wall94extends radially inward towards the inlet inner wall92as the STS inlet66extends axially downstream towards (or to) the combustion chamber56and its outer wall64. The inlet outer wall94ofFIG.10may thereby be canted relative to the axial centerline22, and converge radially inwards towards the inlet inner wall92. With this configuration, a cross-sectional flow area of the core flowpath46(e.g., when viewed in a plane perpendicular to the axial centerline22) may continuously (or intermittently) decrease as the STS inlet66extends downstream axially along the axial centerline22towards (or to) the combustion chamber56. Of course, in other embodiments, the inlet inner wall92may converge radially outwards toward the inlet outer wall94and the inlet outer wall94may maintain a substantially constant radius, or vice versa.

Referring toFIG.11, the STS inlet66may be configured to diffuse and decelerate the exhaust gas directed towards/into the combustion chamber56. The STS inlet66may thereby increase a static pressure of the exhaust gas provided to the rotating detonation combustor52. The STS inlet66ofFIG.11, for example, is configured as a convergent-divergent inlet nozzle (e.g., diffuser inlet), which inlet nozzle is axially between and fluidly couples the turbine section30and the combustion chamber56.

The STS inlet66ofFIG.11includes the radial inner wall92and the radial outer wall94. The inlet inner wall92extends circumferentially about (e.g., completely around) the axial centerline22. An upstream portion96A of the inlet inner wall92extends radially outward towards the inlet outer wall94as the STS inlet66extends axially downstream, towards the combustion chamber56, to an inner wall peak98. The inlet inner wall upstream portion96A ofFIG.11may thereby be canted relative to the axial centerline22, and converge radially outward towards the inlet outer wall94. A downstream portion96B of the inlet inner wall92extends radially inwards away from the inlet outer wall94as the STS inlet66extends axially downstream away from the inner wall peak98and towards (or to) the combustion chamber56and its inner wall62. The inlet inner wall downstream portion96B ofFIG.11may thereby be canted relative to the axial centerline22, and diverge radially inward away from the inlet outer wall94. The inlet outer wall94extends circumferentially about (e.g., completely around) the axial centerline22. The inlet outer wall94extends radially inward towards the inlet inner wall92as the STS inlet66extends axially downstream towards (or to) the combustion chamber56and its outer wall64. The inlet outer wall94ofFIG.10may thereby be canted relative to the axial centerline22, and converge radially inwards towards the inlet inner wall92. With this configuration, a cross-sectional flow area of the core flowpath46(e.g., when viewed in a plane perpendicular to the axial centerline22) may continuously (or intermittently) decrease as the STS inlet66extends downstream axially along the axial centerline22to the inner wall peak98. The cross-sectional flow area of the core flowpath46may continuously (or intermittently) increase as the STS inlet66extends downstream axially along the axial centerline22away from the inner wall peak98and towards (or to) the combustion chamber56. Of course, in other embodiments, the inlet inner wall92may converge and then diverge and the inlet outer wall94may maintain a substantially constant radius, or otherwise.

Referring toFIGS.12,13,14A and14B, the turbine engine20may be configured with a flow regulator100. This flow regulator100may be configured to regulate flow of the exhaust gas into the supplemental thrust section31and its combustion chamber56. Various embodiments of the flow regulator100are described below with respect toFIGS.12,13,14A and14B. While these embodiments are described separately, any two or more of these embodiments may be combined together to provide further flow control. By way of example, the embodiment ofFIG.12or13may be paired with the embodiment ofFIG.14A or14B. In another example, the embodiment ofFIG.12may be paired with the embodiment ofFIG.13and the embodiment ofFIG.14A or14B. The flow regulator100of the present disclosure, however, is not limited to the exemplary flow configurations described herein.

Referring toFIG.12, the flow regulator100may be configured as a variable area diffuser102. The inlet inner wall92ofFIG.12, for example, includes a movable sleeve104; e.g., a translating sleeve. This movable sleeve104is configured to move (e.g., translate) axially along the axial centerline22between a closed (e.g., aft, downstream) position (see solid line sleeve104) and an open (e.g., forward, upstream) position (see dashed line sleeve104). This movable sleeve104is configured to form/carry the inner wall peak98and at least adjacent portions (or the entirety of) the upstream and downstream portions96A and98B of the inlet inner wall92. With this configuration, the movable sleeve104is operable to move the inner wall peak98axially between an aft, downstream position when closed and to a forward, upstream position when open. At the aft, downstream (e.g., closed) position, the inner wall peak98is disposed a radial first distance106A (e.g., measured along a line perpendicular to the axial centerline22) from the inlet outer wall94. At the forward, upstream (e.g., open) position, the inner wall peak98is disposed a radial second distance106B (e.g., measured along a line perpendicular to the axial centerline22) from the inlet outer wall94. The second distance106B is greater than the first distance106A. Thus, the cross-sectional area of the STS inlet66is greater when the movable sleeve104is in its open position than when the movable sleeve104is in its closed position. Of course, although the movable sleeve104is configured as part of the inlet inner wall92inFIG.12, the inlet outer wall94may also or alternatively be configured with a movable sleeve104for regulating flow through the STS inlet66; e.g., see alsoFIG.18A.

Referring toFIG.13, the flow regulator100may be configured as a variable area vane array108. This vane array108may be configured at, upstream of or downstream of the STS inlet66(seeFIGS.10and11). The vane array108ofFIG.13includes a plurality of variable stator vanes110arranged circumferentially about the axial centerline22in an annular array. Each of these variable stator vanes110extends radially across the core flowpath46between an inner duct wall112(e.g., the inlet inner wall92or an adjacent wall) and an outer duct wall114(e.g., the inlet outer wall94or an adjacent wall). Each of the variable stator vanes110is configured to pivot about a (e.g., radially extending) pivot axis115between an open position (see solid line vane) and a closed position (see dashed line vane). When the variable stator vanes110are in their open positions, a cross-sectional flow area between each adjacent set of vanes110is greater than the cross-sectional flow area when the variable stator vanes110are in their closed positions. The variable stator vanes110may also or alternatively be configured to impart or remove swirl from the incoming exhaust gas.

Referring toFIGS.14A and14B, the flow regulator100may be configured as a valve116. The valve116ofFIGS.14A and14Bis configured to open and close or otherwise regulate flow through a supplemental thrust section bypass passage118configured to bypass at least a portion of entirety of the supplemental thrust section31. The valve116, for example, may be arranged at (e.g., on, adjacent or proximate) a bleed orifice120of the supplemental thrust section bypass passage118to the core flowpath46. When the valve116is closed, the exhaust gas may flow unimpeded through the duct90to the rotating detonation combustor52and its combustion chamber56. However, when the valve116is open, at least some of the exhaust gas may be bleed (or otherwise diverted) from the core flowpath46into the supplemental thrust section bypass passage118.

In some embodiments, referring toFIG.14A, the supplemental thrust section bypass passage118may be configured to rejoin the core flowpath46downstream of the supplemental thrust section31. In other embodiments, referring toFIG.14B, the supplemental thrust section bypass passage118may be configured to remain separate from the core flowpath46. The supplemental thrust section bypass passage118and the core flowpath46, for example, may discretely direct the exhaust gas out of the turbine engine20.

While the turbine engine20of the present disclosure is configured with the rotating detonation combustor52and, more generally, the supplemental thrust section31to facilitate high thrust operation as described above, the turbine engine20may spend more time during its life cycle at low thrust operation. To improve low thrust operation efficiency, the valve116ofFIGS.14A and14Bmay be configured to partially or completely open during low thrust operation. The exhaust gas may thereby flow through both the supplemental thrust section31and the supplemental thrust section bypass passage118to reduce flow restriction to the exhaust gas before reaching, for example, the exhaust nozzle38; seeFIG.1. The valve116may also be partially opened during low power supplemental thrust section operation. Of course, during other modes of operation, the valve116may completely close off the supplemental thrust section bypass passage118during supplemental thrust section operation.

Referring toFIG.15, the turbine engine20may also or alternatively be configured with a hot-section bypass passage122. This hot-section bypass passage122is configured to bleed core gas from upstream of at least the combustor section29, and direct that bled core air for use in the combustion chamber56. The hot-section bypass passage122ofFIG.15, for example, extends between and to a bypass passage inlet124and a bypass passage outlet126. The bypass passage inlet124ofFIG.15is located at (e.g., on, adjacent or proximate) the compressor section28, and is configured to bleed the core air from the core flowpath46within (or upstream of, or downstream of) the compressor section28. The bypass passage outlet126ofFIG.15is located at (e.g., on, adjacent or proximate) the rotating detonation combustor52and its combustion chamber56, and is configured to direct the bled core air to the rotating detonation combustor52. With this configuration, the hot-section bypass passage122may direct a portion of the core air to bypass the combustor section29and the turbine section30in order to provide the supplemental thrust section31with a supply of relatively cool, low speed and oxygen rich gas during supplemental thrust section operation. This additional air may facilitate improved and/or more stable rotating detonation combustor operation.

The flow of bled core air through the hot-section bypass passage122may be regulated by a valve128. The valve128may be completely opened during one or more modes of supplemental thrust section operation. The valve128may be partially opened during one or more other modes of supplemental thrust section operation. The valve128may be closed while the supplemental thrust section31is non-operational, or during low power supplemental thrust section operation.

In some embodiments, referring toFIG.16A, the exhaust nozzle38may be configured as a fixed exhaust nozzle130. In other embodiments, referring toFIG.16B, the exhaust nozzle38may be configured as a variable exhaust nozzle132. The variable exhaust nozzle132is configured to change a cross-sectional flow area of the exhaust nozzle to provide enhanced flow and/or pressure control during, inter alia, supplemental thrust section operation.

In some embodiments, referring toFIG.17A, the exhaust nozzle38and its outlet26may have an annular cross-sectional configuration. In other embodiments, referring to FIG.17B, the exhaust nozzle38and its outlet26may have a solid (e.g., non-annular, cylindrical, etc.) cross-sectional configuration.

In some embodiments, the bypass supplemental thrust section bypass passage118(e.g., seeFIGS.13A and13B) may be configured as a single tubular conduit, or include a set of tubular conduits.

In some embodiments, referring toFIGS.18A-18D, the bypass supplemental thrust section bypass passage118may include one or more annular ducts134A and134B (generally referred to as “134”). The bypass supplemental thrust section bypass passage118ofFIGS.18-A-18D, for example, includes the annular inner duct134A and the annular outer duct134B. The inner duct134A is arranged inward of the rotating detonation combustor52, and extends axially along the inner walls62and68. The rotating detonation combustor52thereby circumscribes the inner duct134A. The outer duct134B circumscribes the rotating detonation combustor52, and extends axially along the outer walls64and70. The rotating detonation combustor52is thereby arranged radially inward of the outer duct134B. The ducts134are configured for routing bypass flow around the rotating detonation combustor52or the entire supplemental thrust section. One or more of the ducts134may also or alternatively be configured for cooling a respective sidewall (e.g.,62and/or68;64and/or70) of the rotating detonation combustor52.

Referring toFIG.18A, the inner duct134A has a radial flowpath dimension136A. The outer duct134B has a radial flowpath dimension136B. The rotating detonation combustor52has a radial flowpath dimension138, which may be equal to or different (e.g., greater or less) than the radial flowpath dimension136A and/or136B.

Referring toFIG.18A-18C, bypass flow through one or each of the ducts134may be regulated by a respective flow regulator100. In some embodiments, referring toFIG.18A, each flow regulator100may be configured as a moveable sleeve140A,140B. In some embodiments, referring toFIG.18B, each flow regulator100may be include a plurality of variable stator vanes142A,142B, or a rotating blocker door. In some embodiments, referring toFIG.18C, at least one of the ducts (e.g.,134B) may be configured without a flow regulator.

Referring toFIGS.18B and18D, flow into the rotating detonation combustor52may also or alternatively be regulated by a respective flow regulator100. In some embodiments, referring toFIG.18B, the flow regulator100may be include a plurality of variable stator vanes144, or a rotating blocker door. In some embodiments, referring toFIG.18D, the flow regulator100may be configured as a translating plug146.

In some embodiments, the flow regulator(s)100may fully close the respective duct(s)134when the rotating detonation combustor52is operational. In other embodiments, the flow regulator(s)100may partially close the respective duct(s)134when the rotating detonation combustor52is operational. This may facilitate, for example, low power rotating detonation combustor operation, cooling of the rotating detonation combustor sidewalls during rotating detonation combustor operation, providing a mixed flow exhaust stream aft of the rotating detonation combustor52, etc. Furthermore, while the flow regulator(s)100may be fully open when the rotating detonation combustor52is non-operational, the flow regulator(s) may also be partially closed during certain modes of engine operation while the rotating detonation combustor52is non-operational.

In some embodiments, referring toFIG.1, the combustor section29may be configured as an axial flow combustor section. In other embodiments, referring toFIG.19, the combustor section29may be configured as a reverse flow combustor section.

The supplemental thrust section31and its rotating detonation combustor52may be included in various turbine engines other than the ones described above. The supplemental thrust section31, for example, may be included in a geared turbine engine where a gear train connects one or more shafts to one or more rotors in a compressor section and/or any other engine section. Alternatively, the supplemental thrust section31may be included in a turbine engine configured without a gear train. The supplemental thrust section31may be included in a geared or non-geared turbine engine configured with a single spool (seeFIGS.1and19), with two spools, or with more than two spools. The present disclosure therefore is not limited to any particular types or configurations of turbine engines.