Abstract:
A turbine engine has a circumferential array of combustion chamber conduits downstream of the compressor and upstream of the turbine. Means are provided for directing oxygen-containing gas from the compressor to the conduits so as to cyclically feed a gas charge into each conduit through its first port and permit discharge of combustion products of the charge and fuel through both the first and second ports.

Description:
BACKGROUND OF THE INVENTION 
   (1) Field of the Invention 
   This invention relates to turbine engines, and more particularly to hybrid pulse combustion turbine engines. 
   (2) Description of the Related Art 
   Diverse pulsed combustion technologies exist. Pulse detonation engines (PDE&#39;s) represent areas of particular development. In a generalized PDE, fuel and oxidizer (e.g. oxygen-containing gas such as air) are admitted to an elongate combustion chamber at an upstream inlet end, typically through an inlet valve as a mixture. Upon introduction of this charge, the valve is closed and an igniter is utilized to detonate the charge (either directly or through a deflagration to detonation transition). A detonation wave propagates toward the outlet at supersonic speed causing substantial combustion of the fuel/air mixture before the mixture can be substantially driven from the outlet. The result of the combustion is to rapidly elevate pressure within the chamber before substantial gas can escape inertially through the outlet. The effect of this inertial confinement is to produce near constant volume combustion as distinguished, for example, from constant pressure combustion. 
   PDE technology has a variety of applications. A traditional application is pulsejet engines. Certain recent applications involve use in turbine or hybrid engines. U.S. Pat. No. 6,442,930 and earlier patents identify several hybrid applications. These include uses as thrust augmentors and as replacements for conventional continuous constant pressure gas turbine combustors. 
   BRIEF SUMMARY OF THE INVENTION 
   Accordingly, one aspect of the invention relates to a turbine engine having a circumferential array of combustion chamber conduits downstream of a compressor section and upstream of a turbine section. The engine has means for directing oxygen-containing gas from the compressor to the conduits so as to cyclically feed a charge into each of the conduits through a first port and permit discharge of products of combustion of the charge and fuel through both the first port and a second port. The engine may further include means for directing the products from the ports to the turbine and mixing the products with a flow from the compressor bypassing the conduits so as to present the turbine section with a circumferentially uniform flow. The conduits may rotate about an engine axis. The directing means may have at least a first portion non-rotating about the axis. The turbine and compressor each may have high and low stages on respective high and low spools and the conduits may be on a free spool. A final stage of the compressor may also be on the free spool. Rotation of the conduits may be driven by partially tangential direction of the products. There may be a first airflow through the compressor and turbine, a first portion of which passes through the combustion chamber conduits in the charges. A second portion may bypass combustion. A mass flow ratio of the first portion to the second portion may be between 1:1 and 1:3. The engine may be a turbofan and the first airflow may be a core airflow, with a bypass airflow bypassing the compressor and turbine. The combustion may comprise detonation. The engine may have a number of igniters, each of which is positioned relative to an associated one of the conduits to ignite the combustion of the charge in such conduit. 
   Another aspect of the invention relates to a turbofan engine. A turbine, coaxial with a compressor, drives the compressor and a fan. A pulsed combustion combustor receives air from the compressor and outputs combustion gases to the turbine. A number of combustion chamber conduits each extend along a length between first and second ends and have an outboard portion proximate the first end and an inboard portion proximate the second end. The conduits are held for rotation about the axis through a number of positions. In at least one charge-receiving position, the conduits receive a charge from upstream. In at least one initiation position, combustion of the charge is initiated. In at least one discharge position, products of the combustion are discharged. The charge may be received through the outboard portion and may partially pass therefrom into the inboard portion. The combustion products may be discharged partially through the first end and partially through the second end. The inboard portion may have a partially tangential orientation at the second end. There may be at least one fuel injector for injecting fuel into air from the compressor to form the charges. There may be at least one ring of foils rotating with the conduits as a unit. 
   Another aspect of the invention relates to a pulsed combustion device having a first means for receiving an airflow moving at least partially in a first axial direction and redirecting the airflow to move at least partially in an opposite second axial direction. A combustor assembly has a number of combustion conduits in a circumferential array. The array is rotatable about the axis relative to at least a portion of the first means. Each of the conduits has a first port and a second port. The first port cyclically receives a charge of the airflow and a fuel and the first and second ports cyclically discharge combustion products of the charge. At least one ignition means is positioned to initiate the combustion. The device may be located in a turbine engine having a compressor upstream of the device and a turbine downstream of the device. 
   The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects, and advantages of the invention will be apparent from the description and drawings, and from the claims. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a partial longitudinal sectional view of a turbofan engine. 
       FIG. 2  is a partial longitudinal sectional view of a combustor the engine of  FIG. 1  along a charging sector. 
       FIG. 3  is a radially inward view of a combustion tube of the combustor of FIG.  2 . 
       FIG. 4  is a view of the combustor of FIG.  2 . 
       FIG. 5  is a partial longitudinal sectional view of the combustor of the engine of  FIG. 1  along a discharging sector. 
       FIG. 6  is a partial sectional view of the engine of  FIG. 1  taken along line  6 — 6 . 
       FIG. 7  is a partial sectional view of the engine of  FIG. 1  taken along line  7 — 7 . 
       FIG. 8  is an isolated view of a combustor and mixing duct subassembly of the engine of FIG.  1 . 
       FIG. 9  is a view of the subassembly of  FIG. 8  with combustor core walls removed to show detail. 
       FIG. 10  is an exploded view of the subassembly of FIG.  9 . 
   

   Like reference numbers and designations in the various drawings indicate like elements. 
   DETAILED DESCRIPTION 
     FIG. 1  shows a turbofan engine  20  having a duct  22  and a core  24 . The duct is supported relative to the core by vanes  26 . Of inlet air entering the duct, a fan  28  drives a bypass portion along a first flowpath radially between the duct and the core and core portion along a second flowpath through the core. In the core downstream of the fan, a compressor section  30  having alternating rings of rotor blades and stator vanes compresses the core air and delivers it further downstream to a combustor section  32  where it is mixed with fuel and combusted. A combustion bypass portion of the core air may bypass the combustion within the combustor and be mixed with the combustion portion in a mixing duct  34  downstream of the combustor. Downstream of the mixing duct, a turbine section  36  is driven by the mixing duct output to in turn drive the compressor and fan. An augmentor (not shown) may be located downstream of the turbine. 
   The exemplary engine has three spools: high and low speed spools; and a free spool. The high speed spool includes a high speed shaft  38  and disks and blades of a high speed (and pressure) portion  40  of the compressor section  30  and of a high speed (and pressure) portion  42  of the turbine section  36 . The low spool includes a low speed shaft  44  and the disks and blades of low speed (and pressure) portions  46  and  48  of the compressor and turbine sections. The spools rotate coaxially about the engine central longitudinal axis or centerline  500 . The high compressor portion  40  is driven by the high turbine portion  42 . The fan  28  and low compressor portion  46  are driven by the low turbine portion  48 . 
   The free spool comprises a circumferential array of combustor tubes  50  (FIG.  2 ). The tubes are held by a pair of fore and aft disks  52  and  54 . At a forward end of a web  56  extending forward from the fore disk  52 , a ball bearing system  58  connects the free spool to the high spool shaft. Similarly, at the aft end of a web  60  extending aft from the aft disk  54 , a roller bearing system  62  connects the free spool to the high spool shaft  38 . The free spool further includes a disk  70  carrying the last stage blades  72  of the compressor section  30 . In the exemplary embodiment, the three disks of the free spool are separately formed (e.g., via casting and machining) then securely fastened to each other (e.g., via bolting). 
   Each exemplary tube  50  is formed as a generally u-shaped tube of circular cross-section having a centerline  510  and extending from a first end  80  to a second end  82 . A first leg  84  of the tube extends from the first end  80  forward to a transition  86 . A second leg  88  extends forward from the second end  82  to the transition  86 . In the exemplary embodiment, the first leg  84  is straight and extends parallel to the axis  500 . The first leg  84  is secured within radially outboard apertures  90  and  92  of the fore and aft disks  52  and  54 . An exemplary transition  86  is slightly more than a semitoroid, extending radially from a forward end of the first leg  84  inward to a forward end of the second leg  88 . In the exemplary embodiment, the major portion of second leg  88  extends approximately coplanar to the first leg and the axis  500  but at a slight angle θ 1  thereto. The second leg is secured within apertures  94  and  96  of the fore and aft disks  52  and  54 . An aft portion  100  of the second leg  88  is directed at least partially tangentially. In the exemplary embodiment, the aft portion  100  is nearly skew to the axis  500  and local longitudinal centerplane of the engine by an angle θ 2  (FIG.  3 ). An igniter  102  is mounted to the tube. An exemplary mounting location for a single igniter is at the apex or forwardmost portion of the transition  86 . 
   In the exemplary embodiment, a core flowpath ( FIG. 2 ) from the compressor has two main portions: an inboard portion  104  in an inboard duct  105 ; and an outboard portion  106  in an outboard duct  107 . In the exemplary embodiment, the outboard flowpath  106 /duct  107  is bounded on an outboard side by an inboard surface of a non-rotating outer core wall  108 . The paths  104  and  106  and their associated ducts are separated by a non-rotating intermediate wall  110  running generally parallel to the outer wall  108 . In the exemplary embodiment, the inboard flowpath  104 /duct  105  is bounded on its inboard side by a combination of an outboard duct wall  112  extending aft from an outboard platform portion of the disk  70  to a platform stub  114  of the disk  52  where it merges with the web  56 . Further downstream, the flowpath  104 /duct  105  is bounded by outboard surfaces of respective aft and fore platforms  116  and  118  of the disks  52  and  54 . The disk  54  has an aft platform  120  having a shoulder interfitting with a fixed inner wall  122  of the mixing duct  34  to continue the flowpath  104  into the mixing duct  34 . This interfitting permits relative rotation. 
   The intermediate wall  110  serves to divide the core flow into the paths  104  and  106 . Along charging sectors of the circumference of the core, the outer wall  108  and intermediate wall  110  diverge generally radially outward from fore-to-aft and upstream-to-downstream until approximately the longitudinal location of the tube first ends  80 . Downstream thereof, the walls  108  and  110  turn radially inward and eventually back forward at respective aft portions  130  and  132 . In the exemplary embodiment, the outer and intermediate walls  108  and  110  extend further inward and forward to meet the tube first ends  80 , with interior surfaces of the walls being approximately aligned with interior surfaces of the tubes. End portions  134  and  136  of the walls  108  and  110  diverge respectively inward and outward to accommodate a ring  138  having apertures  140  into which the first end portions of each tube are secured. The exemplary ring has no other apertures. 
   A plurality of fuel injectors  142  are secured in the outer core wall  108  to introduce fuel  144  into the combination portion of core air flowing along/in the outboard flowpath  106 /duct  107 . With the tubes in the charging sectors, the combustion bypass portion of core air flowing along/in the inboard flowpath  104 /duct  105  passes around the tubes  50  through associated apertures  145  ( FIG. 4 ) in the disks  52  and  54  between the tubes. That portion passes around the tube second end  82  into the mixing duct  34  between its inboard wall  122  and outboard wall  146 . Along this charging sector, the inboard and outboard walls  122  and  146  tend to be relatively close to each other at their forward ends, with the inboard wall  122  being slightly inboard of the tube second ends  82  and the outboard wall being approximately radially between the tube ends  80  and  82 . Along this charging sector, the combustion portion of the core air passes along in the outboard flowpath  106 /duct  107 , mixing with the fuel  144  at an upstream end of the flowpath  106 /duct  107 . This flowpath extends generally rearward, then turns inward and forward in a downstream flowpath area  148  between the aft portions  130  and  132  of the walls  108  and  110 . This flow portion (including/carrying the injected fuel) then proceeds through the tube first ends  80 , forwardly through the first legs  84 , through the transitions  86 , and then through the second legs  88  toward the second ends  82 . 
   As the free spool rotates, each tube transiently progresses through a charging sector to a subsequent discharging sector.  FIG. 5  shows a tube  50  in a discharging sector. In the exemplary implementation, along the discharging sectors the flow is still divided into inboard and outboard flowpaths  104  and  106  bounded by generally the same surfaces of the same elements as in the charging sector of FIG.  2 . In the discharging sector, however, the wall aft portions  130  and  132  and thus the outboard duct  107  terminate proximate an outboard extreme of the adjacent tube first ends  80  so that the outer core flowpath  106 /duct  107  is locally blocked by the outboard surface of a forward continuation  150  of the mixing duct outboard wall  146 . Along this discharging sector, the front end of the outboard wall  146  is relatively far apart from the adjacent portion of the inboard wall  122 . Thus along this sector, both tube ends  80  and  82  are open to the interior of the mixing duct. The local blocking of the flowpath  106 /duct  107  causes flow therethrough to be directed generally tangentially toward an adjacent charging sector. In the exemplary implementation, along the discharging sectors there are no fuel injectors. The inboard flowpath  104  may proceed as in the charging sector. 
     FIGS. 8-10  show further details of the exemplary combustor  32  and mixing duct  34 . The fuel injectors  142  are shown mounted in the outer core wall  108  in an exemplary three groups of five injectors. The injectors have bodies extending outboard from the wall  108  with inlets coupled to a single common fuel rail  170  for delivering fuel to the injectors. Injector outlets may be flush, subflush, or slightly protruding from the inboard surface of the wall  108 . Extending slightly radially inward and aft from the inboard surface of the wall  108  are a plurality of guide walls  172 . The exemplary guide walls are positioned between adjacent injectors of each group and at the outboard extremes of the groups. The guide walls serve to encourage uniform mixing of the fuel and air. 
   As a tube  50  rotates to the beginning of each discharging sector, its igniter  102  is triggered to ignite the fuel/air mixture within the tube  50 . Exemplary igniters may take the form of sparkplugs rotating past magneto points (not shown) or otherwise having a wiring harness (not shown) coupling the igniters to electrical contacts spanning the rotating and non-rotating portions of the engine (e.g., as in various distributor constructions and techniques). 
   The flame front resulting from the ignition simultaneously propagates rearward down both legs of the tube causing combustion products to be expelled from both the ends  80  and  82 . The expulsion from both ends of a single tube may simplify or eliminate sealing considerations (e.g., as compared with the situation where a valve structure seals an inlet end of a tube so that combustion products are exclusively expelled from an outlet end). 
   The angling of the aft portion  100  produces tangential thrust to drive the rotation of the free spool. In the exemplary embodiment, this applies sufficient torque to rotate the free spool at a desired rotational speed. In an exemplary three spool engine, an exemplary steady state rotational speed of the carousel is 1500-4500 RPM. The specific operating range will be influenced by engine dimensional considerations in view of carousel structural integrity and the number of charge/discharge cycles per rotation. A narrower range of 2500-3500 target RPM is likely with the lower third of this range more likely for the exemplary three cycle/rotation engine and the upper third for a one cycle/rotation engine. In operation, these speeds will likely be substantially lower than the high spool speed and approximately the same or moderately lower than the low spool speed. Other variations and embellishments may relate to or control rotation of the free spool. These may include the presence of additional rotating blades or fixed vanes. An initial rotation may be provided by the engine starter motor (not shown) or by a dedicated starter motor (not shown) for the combustor. 
   The combustion products discharged from the tube ends  80  and  82  mix with the air in the flowpath  104  at the front end of the mixing duct  30 , mixing more fully as the combined flow moves aft/downstream. The mixing duct outer wall  146 , which locally forms a portion of the core outer wall, has radially elevated portions or humps  160  at the duct front end aligned with the discharge sectors and intervening inboard portions  162  at the duct front end aligned with the charging sectors. Progressing rearward from the front end of the mixing duct, both the inboard and outboard walls  122  and  126  diverge generally radially outwardly to join inner and outer core walls at the high turbine  42 . During this transition, however, the elevated portions  160  broaden until the outer wall  146  is essentially annular at the aft end of the mixing duct. The humps also are disposed helically in accordance with the normal operating tangential velocity component of gases exiting the combustor. Presenting a relatively uniform mixing duct cross section to the face of the high pressure high turbine, the mixing duct acts as a pulse dampener to smooth out the flow presented to the turbine. 
   The magnitude of the pressure pulses from the individual combustion tubes is minimized by the apparent high firing frequency (e.g., 1000 Hz to 6000 Hz) provided by the rotating tube pack. The pressure pulses may overlap at a relatively uniform peak level resulting in a quasi-steady state effluent pressure. Additional viscous dampening of any residual small cyclical pressure variations may occur in the mixing duct. 
   In the exemplary embodiment, the combustion tubes  50  may be operated in a detonation mode. In such a mode, when triggered, each igniter produces a detonation pulse which propagates a flame front radially outward from an associated ignition point at a supersonic speed (e.g., over about 3,000 feet per second (fps) any typically in the range of 4,000-6,000 fps). Near total combustion will be achieved in the time required for the flame front to travel to the tube ends (half the total tube length). In an exemplary detonation operation, the operating pressure ratio (OPR) of the combustion tubes may generally be between two and six. 
   Alternative embodiments might operate in a deflagration mode. This might be achieved by using a multi-igniter, multi-point ignition system. Various embellishments may be made involving the positioning of the igniters and their trigger timing to achieve desired combustion parameters. In addition to discrete multi-point ignition, there may be continuous (e.g., laser or ion or other energy beam) or multi-continuous ignition. In such a deflagration mode, with a sufficiently high number of ignition points (and a sufficiently small spacing therebetween) substantial combustion can be achieved through subsonic deflagration in a sufficiently short time that may approach or even be shorter than that of the detonation system. With a short enough combustion time, the inertial confinement of the gas in the tube (the tube ends both being open at the discharge sector) is effective to achieve near constant volume combustion. The resulting effect can be achievement of efficiencies similar to the detonation system without some or all of the attendant drawbacks (e.g., narrow operating parameters, supplemental oxygen, exotic chamber geometries, noise, and the like). In a deflagration mode, the flame fronts propagate outwards from their ignition point(s) at a subsonic speed (e.g., under about 2,000 feet per second (fps) any typically in the vicinity of 1,000 fps). Hypothetical deflagration devices may operate in a broad performance envelope. The exemplary OPR&#39;s may be between 1.5 and five. The low end of the spectrum might be typical for ram-fed applications and the high end for pre-compression applications. Basic applications include use as pulsejet engines. In an exemplary ram-fed pulsejet engine, inlet air is fed into the device due to the movement of the vehicle through the air and the exhaust expelled from the outlet as thrust. 
   In exemplary embodiments, there may be between ten and fifty combustion tubes, more narrowly, twenty and forty. Exemplary tube lengths (and the approximately similar combustion chamber lengths) are between forty cm and one m, more narrowly, sixty cm and eighty cm. The exemplary tube cross sectional areas are between eight cm 2  and seventy cm 2 , more narrowly, fifteen cm 2  and fifty cm 2 . The exemplary engine has three pairs of charging and discharging sectors so that each combustion tube undergoes three charge/discharge cycles in one rotational cycle. Other numbers of pairs are possible (e.g., 1-10 pairs). The number of mixing duct humps or other related features would vary correspondingly. Each charging sector may be larger than its associated discharging sector. In the exemplary three cycle embodiment, an exemplary charging sector is approximately 86° and an exemplary discharging sector is approximately 33°. The key limitation regarding size of the charging sector is the time required to charge the combustion tubes at a given radius from the engine axis and rotational speed. 
   One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. For example, one possible variation is to have the tubes nonrotating but have a rotating manifold structure. Various additional structure may be added, such as having turning vanes or a turbine stage on the free spool. Additionally, the principles may be applied to various turbine applications and to nonturbine applications. An exemplary nonturbine application is used as a rotating pulse jet. Accordingly, other embodiments are within the scope of the following claims.