Abstract:
A turbine engine has a case with an axis. A fan is mounted for rotation about the axis. A turbine is mechanically coupled to the fan to drive rotation of the fan about the axis. A number of compressor/turbine units are downstream of the fan and upstream of the turbine along a core flowpath. A number of compressors are coupled to the compressor/turbine units to receive air and deliver combustion gas to drive the turbine.

Description:
U.S. GOVERNMENT RIGHTS 
   The invention was made with U.S. Government support under contract F33615-95-C-2503 awarded by the U.S. Air Force. The U.S. Government has certain rights in the invention. 

   BACKGROUND OF THE INVENTION 
   This invention relates to engines, and more particularly to hybrid pulse combustion turbine engines. 
   In a conventional gas turbine engine, combustion occurs in a continuous, near constant pressure (Brayton cycle), mode. Although present gas turbine engine combustors are relatively efficient, the thermodynamic benefit to cycle efficiency associated with performing the combustion operation at a higher time-averaged pressure has led to many efforts to improve combustion. 
   It has been proposed to improve thermodynamic efficiency by applying the more efficient combustion of near constant volume combustion pulse detonation engines (PDEs) to turbine engine combustors. 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. 
   U.S. Pat. No. 6,442,930, for example, suggests combustor use of PDE technology in addition to use as a thrust augmentor in engines with conventional combustors. Other pulsed combustors are shown in U.S. Pat. Nos. 6,886,325 and 6,901,738. 
   BRIEF SUMMARY OF THE INVENTION 
   One aspect of the invention involves a turbine engine having a case with an axis. A fan is mounted for rotation about the axis. A turbine is mechanically coupled to the fan to drive rotation of the fan about the axis. A number of compressor/turbine units are downstream of the fan and upstream of the turbine along a core flowpath. A number of compressors are coupled to the compressor/turbine units to receive air and deliver combustion gas to drive the turbine. 
   In various implementations, the compressor/turbine units may be centrifugal compressor/radial turbine units, with the turbine coaxially driving the impeller by means of a connecting shaft. There may be a circumferential array of the compressor/turbine units and a circumferential array of the combustors. Each of the compressor/turbine units may be uniquely associated with a single one of the combustors and vice versa. The compressor/turbine units may be coupled to the combustor so that: the compressor of the compressor/turbine unit delivers air to the associated combustor; and the turbine of the compressor/turbine unit receives the combustion gas from the associated combustor. The turbine may be an axial turbine receiving the combustion gas from all of the compressor/turbine units. The axial turbine may be co-spooled with the fan. There may be at least eight of the compressor/turbine units and at least eight of the combustors. The combustors may be non-rotating. 
   Another aspect of the invention involves a method for operating a turbine engine. Air is directed from a fan to a number of compressor/turbine units. The air is compressed in the compressor/turbine units. The air is directed to a number of combustors. The air is combusted with fuel in the combustors to produce combustion gas. Work is extracted from the combustion gas in the compressor/turbine units to drive the compression. The combustion gas is directed from the compressor/turbine units to a turbine. Work is extracted from the combustion gas in the turbine to drive rotation of the fan. 
   In various implementations, the combustion gas may be directed from the turbine to join a bypass flow of air from the fan. A mass flow ratio of the flow of the air delivered to the combustors to the bypass flow may be between 1.1 and 1:3. The combusting may be a pulse combusting. The combusting may comprise detonation. The combusting may comprise operating respective ones of the combustors out of phase with each other. The method may be used in aircraft propulsion. 
   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 schematic partial longitudinal sectional view of a turbofan engine. 
       FIG. 2  is a cutaway view of the engine of  FIG. 1 . 
       FIG. 3  is a schematic partial longitudinal sectional view of an alternative engine. 
       FIG. 4  is a front schematic view of a second alternative engine. 
   

   Like reference numbers and designations in the various drawings indicate like elements. 
   DETAILED DESCRIPTION 
     FIG. 1  shows a turbofan engine  20  having central longitudinal axis  500 , a case  22 , and a core  24 . The case  22  defines a duct  26  extending from an upstream inlet  28  to a downstream outlet  30 . Of an inlet airflow  510  entering the duct, a fan  32  drives a bypass portion  512  and a core portion  514  along respective bypass and core flowpaths through the duct. The exemplary fan  32  has two blade stages and two interspersed vane stages. The blade stages may be supported on a shaft  34 . 
   As is described in further detail below, the exemplary engine  20  also includes a circumferential array of compressor/turbine units  38 , a combustor section  40  (e.g., circumferential array of combustors  41 ), and a turbine section  42 . Other components (e.g., an augmentor and an exhaust nozzle) may also be present.  FIG. 2  shows further details of exemplary positions of the exemplary compressor/turbine units  38  and combustors  41 . 
   The core airflow  514  is divided by ducts  44  into branching portions directed to the compressor sections  50  (e.g., centrifugal compressors) of each of the units  38 . Rotation of the impeller of the section  50  is driven by the turbine of the turbine section  52  (e.g., a radial turbine) of the associated unit  38 . The units  38  thus compress the flow  514  into compressed flows  516  directed to the combustor section  40 . In each unit  38 , the compressor section  50  and turbine section  52  are coaxial about an axis non-coincident with the engine axis  500 . In the combustor section  40 , the compressed air is mixed with a fuel flow  518  and combusted to form combustion gas  520 . The gas  520  is directed to the turbine of the turbine section  52  where it is partially expanded to extract the work to compress the flow  514 . 
   From the unit  38 , the partially expanded combustion gas flow  522  is directed to the turbine section  42 . For example, the turbine sections  52  of the various units  38  may be coupled to a common discharge manifold  60  feeding an upstream/inlet end of the turbine section  42 . As the flow  522  passes through the turbine section  42  it is further expanded and discharged as a flow  524 . The exemplary flow  524  is directed via a manifold duct  62  to merge with the bypass flow  512  and form a combined flow  526 . This combined flow may ultimately be discharged from the outlet  30 . 
   In the exemplary engine of  FIG. 1 , the blade stages of the turbine section  42  are co-spooled with the fan on the shaft  34 . The positioning of the turbine section  42  forward of the combustor section  40 , along with the generally forward flow through the turbine section  42  facilitates a short shaft  34  and a longitudinally compact engine. The configuration also hides the moving/hot surfaces of the turbine section  42  from line-of-sight exposure through the outlet. This may be advantageous for low observability properties including radar return and infrared signature. 
     FIG. 1  shows further details of the exemplary combustor section  40 .  FIG. 1  shows an inner member  80  within an outer member  82 . The airflow  516  is received through an associated conduit  84  to a volume or space  86  between the inner and outer members. There may be a circumferential array of the inner members  80  (one for each combustor  41 ). In some variations, the outer member  82  may be a single outer member containing all or more than one of the inner members (e.g., an annular outer member). In other variations, there may be a circumferential array of the outer members  82 , each containing an associated one of the inner members  80 . 
   The exemplary inner member  80  has an aft end  90  and a fore end  92 . The exemplary inner member  80  has a first frustoconical wall portion  94  diverging forward from the aft end  90 . The wall portion  94  is foraminate allowing the inflow of air. In the exemplary combustor, a fuel injector  100  may be positioned at the aft end to introduce the fuel flow  518 . An igniter  102  (e.g., a sparkplug) may be positioned to ignite the fuel air mixture to cause combustion. The divergence of the wall portion  94  helps facilitate a deflagration-to-detonation transition. 
   The exemplary inner member  80  has a second wall portion  110  forward of the portion  94 . A convergent wall portion  112  is downstream of the portion  110 . An outlet conduit  114  connects the inner member  80  to the associated turbine section  52 . Individual coupling of the combustors to at least the turbine section  52  prevents crosstalk between the discharge ends of the combustors. This is relevant where the combustors are operated out-of-phase so that the combustion gas discharged by one combustor is not ingested by another. 
   Inlet decoupling is less critical. Thus, there may be a common outer member  82  defining a common inlet plenum. In yet other embodiments, each combustor may be coupled to receive air from the compressor section  50  of one unit  38  while discharging gases to the turbine section  52  of another unit. 
     FIG. 3  shows an alternative configuration with a long shaft  34 ′ connecting a turbine section  42 ′ to the fan. The exemplary turbine section  42 ′ is aft of the combustor section and receives combustion gases from the compressor/turbine unit array through a manifold  160 ′ directing the combustion gases generally aftward and radially inboard of the combustors. The discharged combustion gases and bypass air mix relatively downstream. 
   The effects of the pressure pulses from the individual combustors is minimized by operation out-of-phase with each other. Exemplary firing frequency may be in the vicinity of 50-300 Hz and may vary considerably depending on the scale/size of the engine and resulting impact on combustor section geometry and volume. Various phase combinations are possible, including firing in opposed pairs to limit wobble. Exemplary fan spool speeds are 2000-20000 revolutions per minute (RPM), more narrowly 6000-12000 RPM. Exemplary speeds for the units  38  are 5000-50000 RPM, more narrowly 20000-35000 RPM as an approximation for the 6000-12000 RPM fan spool speeds under steady-state conditions. 
   Many variations are possible. For example, the combustors take a variety of forms, including shapes, positions, and orientations.  FIG. 4  shows an exemplary configuration wherein eight combustors are grouped in two groups concentrated on respective left and right sides of the engine. This creates a wide but small height package which may be advantageous for integration into the airframe of an aircraft (e.g., a fighter aircraft, unmanned aerial vehicle, or missile). 
   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, the details of any particular application will influence the configuration of the combustor. Various features of the combustor may be fully or partially integrated with features of the turbine or the compressor. If applied in a redesign of an existing turbine engine, details of the existing engine may implement details of the implementation. The combustor may alternatively be used in applications beyond turbine engines. Accordingly, other embodiments are within the scope of the following claims.