Patent Publication Number: US-2012023964-A1

Title: Liquid-fueled premixed reverse-flow annular combustor for a gas turbine engine

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This disclosure was made with Government support under N00019-06-C-0081 awarded by The United States Navy. The Government has certain rights in this disclosure. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine and more particularly to a reverse flow annular combustor for an auxiliary power unit (APU). 
     An APU is often utilized to supplement main propulsion engines to provide electrical and/or pneumatic power as well as start the main propulsion engines. APUs are typically a radial or axial gas turbine engine having a compressor, a combustor, and a turbine. The combustor is often a liquid-fueled non-premixed reverse flow annular combustor with an active dome primary combustion zone using liquid fuel injectors to direct a fuel spray into a liner dome section to form a combustible mixture with the air admitted to the dome. 
     SUMMARY 
     A reverse flow annular combustor for a gas turbine engine according to an exemplary aspect of the present disclosure includes a pre-vaporizer/pre-mixing region within a dome section, liquid fuel injectors admitting a fuel spray to that dome section, a combustion region downstream of the pre-vaporizer/pre-mixing region and a dilution region downstream of the combustion region. 
     A method of combustion within a reverse flow annular combustor for a gas turbine engine according to an exemplary aspect of the present disclosure includes: injecting liquid fuel into a liner dome section forming a pre-vaporizer/pre-mixing region within a liner dome section; forming a combustion region downstream of the pre-vaporizer/pre-mixing region; and forming a dilution region downstream of the combustion region. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a partial phantom view of a rotary-wing aircraft illustrating a power plant system; 
         FIG. 2  is a general sectional view of an auxiliary power unit; 
         FIG. 3  is an expanded schematic sectional view of a combustor for the auxiliary power unit falling within the embodiment of the present invention; 
         FIG. 4  is an expanded schematic sectional view of a RELATED ART combustor; 
         FIG. 5A  is a sectional view of a combustor according to one non-limiting embodiment of the present application without the effusion holes shown; and 
         FIG. 5B  is a rear view of the combustor of  FIG. 5A . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a rotary-wing aircraft  10  having a main rotor system  12 . The aircraft  10  includes an airframe  14  having an extending tail  16  which mounts an anti-torque system  18 . The main rotor system  12  is driven about an axis of rotation R through a main rotor gearbox (MGB)  20  by a multi-engine powerplant system  22 —here having three engine packages ENG 1 , ENG 2 , ENG 3  as well as an Auxiliary Power Unit (APU)  24  ( FIG. 2 ). The multi-engine powerplant system  22  generates the power available for flight operations and couples such power to the main rotor assembly  12  through the MGB  20 . Although a particular helicopter configuration is illustrated and described in the disclosed embodiment, other configurations and/or machines, such as high speed compound rotary-wing aircraft with supplemental translational thrust systems, dual contra-rotating, coaxial rotor system aircraft, turbo-props, tilt-rotor, fixed wing aircraft and non-aircraft applications such as ground vehicles will also benefit herefrom. 
     Referring to  FIG. 2 , the APU  24  in the disclosed non-limiting embodiment is a radial gas turbine engine having a turbine wheel  30 T that defines a plurality of turbine blades  34  is disposed opposite a compressor wheel  30 C that defines a plurality of compressor blades  36  about an axis of rotation A. A shaft  38  extends from the turbine wheel  30 T and through the compressor wheel  30 C such that the turbine wheel  30 T and compressor wheel  30 C are coaxially coupled. The compressor blades  36  compress air for communication to a combustor  40  and the turbine blades  34  convert pressure energy of exhaust gases from the combustor  40  into rotational energy. The turbine blades  34  are shaped such that high pressure combusted gases impinge thereon to drive the shaft  38  to thereby convert heat and pressure into mechanical energy. In a similar axial gas turbine engine, the combustion air also exits the combustor radially inwards but is turned axially towards the engine AFT end before entering the axial turbine wheel. 
     With reference to  FIG. 3 , the combustor  40  includes a staged-combustion reverse flow annular combustor design for a radial gas turbine engine. The combustor  40  includes a system of circumferentially spaced liquid fuel injector system (illustrated schematically at I) fueling a pre-vaporizer/pre-mixing region  42 ; a pre-vaporizer/pre-mixing region  42  within a dome section  44  followed by a narrowly controlled combustion region  46  and a downstream dilution region  48 , thereafter having a relatively tight turn  50  into the radially oriented turbine nozzle  34 B and turbine blades  34 . 
     The cross sectional layout of the reverse flow annular combustor  40  follows a typical non-premixed combustor design such as a Rich-Quench-Lean (RQL) combustor, such as proposed within the HSPS/PWA Pyrospin™ RQL Combustor. In the disclosed non-limiting embodiment, the selected fuel injection configuration, size selection and location of OD and ID primary  46 J/dilution  47 /film  43 /dome jets  42 J is such that the liner dome section  44  of the combustor  40  is primarily used for fuel preparation, i.e., atomization, vaporization and mixing while stable combustion is effectively achieved in the relatively short combustion region  46  downstream of the pre-vaporizer/pre-mixing region  42 , i.e., downstream of the OD and ID primary cross-flow air jets  46 J, but before the dilution region  48  and dilution jets  47 . It should be understood that the combustion region  46  may be augmented by, for example, Pyrospin technology, i.e., effusion jet enhanced mixing as indicated by the effusion jets  48  IJ in  FIG. 3 . Flame stabilization in region  46  can be achieved by a properly designed injector body acting (besides its primary function as fuel injector) as material flame holder or by fluidic flame holders given by the cross flow jets  46 J emanating from the primary OD and ID holes. 
     The combustor  40  takes advantage of proven design concepts for traditional liquid-fuel spray non-premixed reverse flow annular combustor designs that have a primary combustion zone located within the liner dome region (RELATED ART;  FIG. 4 ), but employs a liner hole pattern and injector configuration which provides for a pre-vaporizing/pre-mixing type combustion process without the need for a pre-vaporizing system, such as pre-vaporizer tubes, for example. The disclosed staged combustion system and process prevents flash back, reduces complexity and cost associated with a dedicated pre-vaporizing system while preserving the potential benefits of a premixed, pre-vaporized combustion system to provide low NOx emissions, for example. 
     The specific injector system and liner hole size configuration identified as dilution, primary, effusion, film and dome cooling holes for the disclosed staged combustion process provides for the different method of combustion staging—pre-vaporizer/pre-mixing region  42 , combustion region  46  and dilution region  48 —based on a combustor cross sectional configuration originally designed to feature a non-premixed combustion sequence with a primary combustion region located within the liner dome section followed by an intermediate combustion region and a dilution region (RELATED ART;  FIG. 4 ). 
     The conventional combustor (RELATED ART;  FIG. 4 ) includes an air jet arrangement which provide air to both the primary and intermediate combustion regions in which combustion takes place on the fuel rich side in the primary zone, while the intermediate region burns significantly leaner. 
     An air jet arrangement  70  disclosed herein includes pre-vaporizer/pre-mixing region jets  42 J within the liner dome section  44  for the pre-vaporizer/pre-mixing region  42 , primary combustion region jets  46 J, film cooling jets  43  and effusion air jets  48 IJ for the combustion region  46  and downstream dilution jets  47  and effusion air jets  48 J for the dilution region  48  (also illustrated in  FIGS. 3 ,  5 A and  5 B). In order to achieve the prescribed combustor staging, the flow split and the selected hole sizes, numbers and shapes for the air jets defined above will be a function of combustor geometric cross section (including annulus cross section), combustor inlet conditions and injector performance. Since the specific air jet arrangement flow split percentages and hole numbers, sizes and shapes (e.g., tubes, louvers) will depend on the selected combustor geometry or cross section, no specific jet arrangement and flow split need be identified herein. 
     The liner dome section  44  provides the function of a prevaporizer and premixing volume. While, in the present invention, the pre-vaporizer/pre-mixing jets  42 J provide mixing air for the pre-mixing region  42  ( FIG. 3 ), in the related prior art ( FIG. 4 ), the similar jets provide film cooling at  3  and protection of the liner dome section  44  from the combustion products in the primary combustion zone (see  FIG. 4 ). Based on the prescribed reduction of the combustion volume taken by the pre-vaporizer/pre-mixing region  42 , combustion now takes place completely in the region conventionally heretofore utilized as an “intermediate combustion region” (RELATED ART;  FIG. 4 ) with flame holding and stable combustion ensured by the prescribed mechanical or fluidic flame holders. 
     The dilution region  48  disclosed herein provide for enhanced dilution of the hot combustion gases to meet combustor exit flow criteria (such as maximum hot streak temperatures). Those skilled in the art of combustor design will understand that there are likely more than one dilution zone configuration (e.g., with or without transfer tubes), injector system configuration and liner hole size/number/shape configuration to the prescribed conversion from the combustion sequence (RELATED ART  FIG. 4 ) with liner dome primary combustion zone to the sequence illustrated in  FIG. 3  with the liner dome pre-vaporizer/pre-mixing region  42 . 
     After properly designing the air jet arrangement  70  ( FIG. 3 ), no primary air enters the liner dome section  44  as the only air which enters the liner dome section  44  is premix air. That is, the air in the liner dome section  44  is mixing air alone such that an air-fuel mixture that is fuel rich beyond a fuel rich limit so that continuous combustion cannot take place in the liner dome section  44  of the pre-vaporizer/pre-mixing region  42 . 
     Downstream of the pre-vaporizer/pre-mixing region  42  significant dilution of the non-combusting fuel-rich mixture is provided by the primary combustion region jets  46 J and effusion air jets  46 IJ of the air jet arrangement  70  to form the combustion region  46  within which all the combustion takes place. That is, the primary combustion region jets  46 J provide and sustain continuous combustion. While the primary function of the effusion air jets  48 IJ is the protection of the combustor liner walls, they might also support the combustion process via mixing enhancement. Ignition through an igniter arrangement may be achieved in a conventional manner downstream of the liner dome section  44  and within region  42 . However, in the  FIG. 3  configuration, the flame will quickly transition to the intermediate zone at combustion region  46 . Flame stabilization in the combustion region  46  is achieved through the suitable fuel injector arrangement I, and proper injector body design for mechanical and fluidic flame stabilization. The primary air jets  46 J might also contribute to flame stabilization (so called fluidic flame stabilizers) which operate as traverse or cross-flow jets to the internal bulk flow of fuel rich gases travelling from the liner dome section  44  to the turbine nozzle  34 B. 
     Downstream of the combustion region  46 , dilution jets  48 J and effusion air jets  48 IJ are provided for the dilution region  48 . That is, the dilution jets  47  and effusion air jets  48 J (in their secondary function) provide a premixed pre-vaporized combustion system with efficient dilution so as to not damage the turbine blades  34  and turbine nozzle  34 B. 
     It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.